DE112014000391B4 - Low noise and highly efficient aircraft - Google Patents

Abstract

Aircraft with: - an at least partially cylindrically elongated fuselage arrangement (F) with a spatial main extent in the direction of a longitudinal fuselage axis (FA), this fuselage arrangement (F) for accommodating payload as required, the fuselage arrangement (F) at least in sections functioning as a pressurized cabin ( P) can be imprinted with an internal pressure, - a wing assembly (W) firmly attached to the fuselage assembly (F) of the aircraft, for generating lift supporting the aircraft, the force component of the lift generated in horizontal flight perpendicular to the longitudinal axis of the fuselage (FA) opposite the component along the longitudinal axis of the fuselage (FA) clearly predominates, - with a three-point landing gear arrangement (GA) that can be retracted and extended again if necessary, with several rotatably mounted wheels (WE) for statically determined support of the aircraft against the ground level (BO) in floor-level operation, consisting of two structurally the wing arrangement (W) connected main landing gear legs (GAM) and a nose landing gear leg (GAA) structurally connected to the fuselage arrangement (F), whereby the nose landing gear leg (GAA) is arranged along the longitudinal axis of the fuselage (FA) in the direction of flight (FR) in front of the main landing gear legs (GAM), - a motor arrangement ( E), consisting of at least one motor, for generating a drive power available to the aircraft, - at least one rotor (R) bladed with several rotor blades (B) for at least partial conversion of drive power of the motor arrangement (E) into propulsive power to drive the aircraft by rotating the bladed rotor (R) around a geometrical axis of rotation (RA), whereby when the rotor blades (B) of the bladed rotor (R) rotate in rotation, a geometrical rotor surface (RAE) swept over by the rotor blades (B) results, this bladed rotor (R) in its radial spatial extension a smallest inner diameter (ID) and a having the largest outer diameter (OD), this bladed rotor (R) being equipped with blades (B) that are adjustable in the setting angle, this bladed rotor (R) also having a spatial longitudinal extension along its geometric axis of rotation (RA), with its Longitudinal extension of the rotor planes (RP) perpendicular to its geometric axis of rotation (RA) can be identified as a geometric aid to detect a possible area of impact of this rotor (R), - at least one gear (G) for transmitting the drive power of at least one motor of the motor arrangement (E) on at least one bladed rotor (R), - at least one bearing arrangement (S), characterized in that at least one bladed rotor (R) with at least one bearing arrangement (S) is rotatably mounted surrounding the fuselage arrangement (F) and that at least this a bladed rotor (R) approximately in the direction of the longitudinal axis of the fuselage (FA) is axially fixed by at least one bearing arrangement (S) to the fuselage arrangement (F), and that at least this one bladed rotor (R) at the same time surrounds that section of an at least partially cylindrically elongated fuselage arrangement (F) of an aircraft in a ring on the outside in the radial direction , which, viewed in one of the rotor planes (RP) of this rotor (R), can be printed at least partially in the function of a pressure cabin (P) and that at least in one of the rotor planes (RP) of the bladed rotor (R) no further for the essential components of the aircraft are arranged for immediate safe operation of the aircraft, and that at least this one bladed rotor (R) is arranged along the fuselage longitudinal axis (FA), seen in the direction of flight (FR), in front of the wing arrangement (W) and at the same time the geometric axis of rotation ( RA) at least one bladed rotor (R), along a direction of the aircraft vertical axis (VA), starting from the ground plane (BO), viewed in a plane that is immediately upstream of the wing arrangement (W) and perpendicular to the axis of rotation (RA) of this bladed rotor (R), is arranged so that it is above the minimum vertical level resulting in this plane (MV) of the upper surface of the airfoil arrangement (W), is located in such a way that in a rotor plane (RP) of this one bladed rotor (R) the predominant area portion of the geometric area swept over by the rotor blades (B) of the bladed rotor (R) in circulation Rotor surface (RAE), viewed from the ground plane (BO), comes to rest above the wing arrangement (W) in such a way that the ram pressure increase of the propeller jet caused on the wing arrangement (W) downstream against the direction of flight (FR) is at least this one bladed rotor (R) takes place predominantly above the wing arrangement (W) and that at least this one rotor (R) in the direction of the fuselage longitudinal axis (FA) at de The fuselage arrangement (F) between the nose landing gear leg (GAA) and the main landing gear legs of the main landing gear assembly (GAM) is arranged in such a way that, in each of the rotor planes (RP) of this one bladed rotor (R), the largest outside diameter (OD) of this one rotor (R) is smaller than the double minimum distance of the axis of rotation (RA) of this bladed rotor (R) from the ground plane (BO) given by the chassis arrangement (GA), and that at least one motor of the motor device (E) in an area of the aircraft outside the pressurized cabin (P), but at least partially sunk inside a space that flows outwards to the fluid through the outer contour of the wing assembly (W) and the outer skin of the fuselage assembly (F) Aircraft is delimited, and so embedded in this space is arranged that a component that acts as an engine mount structural forces from at least one engine of the The engine assembly (E) discharges in the direction of the structure of the aircraft, is essentially not visible from the outside and that this component in its function of the engine mount outside of the by the outer skin of the wing assembly (W) and from the outer skin of the fuselage assembly (F) formed outer contour essentially no separate, drag-effective surface flushed by the fluid can be assigned, and that at least this one bladed rotor (R) with at least one motor of the motor arrangement (E), which is arranged outside the pressure cabin (P), via gears (F ), is kinematically coupled, and can be set in rotation by this in order to generate a propulsive force that drives the aircraft, the force component of which in the direction of the longitudinal axis of the fuselage (FA) clearly outweighs the force component perpendicular to the longitudinal axis of the fuselage (FA).

Description

The invention relates to an aircraft, preferably an aircraft of the type of commercial aircraft.

Today's aircraft, in particular commercial aircraft, often have a fuselage section that is at least partially cylindrically elongated as part of their fuselage arrangement, the fuselage arrangement being printable, at least in sections, in the function of a pressure cabin. This printable section is printed as a pressure cabin for cruise flight and enables the aircraft to go to greater altitudes with the crew as well as with passengers or cargo, in order to be able to operate there under more advantageous aerodynamic and economic conditions. The overall height of the pressurized cabin within the fuselage arrangement is often such that, for practical and comfort reasons, an upright gait is possible for the average passenger within the pressurized cabin, at least in sections.

In addition, today's commercial aircraft have a landing gear arrangement with several rotatably mounted wheels, with which they can be supported against the ground plane when the aircraft is operating flush with the floor, and which enables the aircraft to take off and land in an essentially horizontal flight direction on elongated runways.

The long-term goal of some interest groups is to further increase the fuel efficiency of transport aircraft, especially commercial aircraft, and to reduce the noise they emit during operation.

So far, an increase in the fuel efficiency of transport aircraft has essentially been achieved through a significant improvement in the propulsion efficiency. On the one hand, this was decisively achieved by replacing turbojet engines with turbojet engines with a bypass flow enveloping the core flow. On the other hand, a steady increase in the bypass flow ratio of these turbofan engines then repeatedly resulted in further improvements in efficiency and thus fuel savings.

According to their architecture, turbofan engines make it possible, with the bypass flow, to guide a large part of the air mass flow flowing through the engine with a comparatively low flow velocity, moved by a fan, around the engine core. The fan, which contributes to the acceleration of most of the mass flow and thus also to a large part of the thrust generation, is driven by the core turbine in the main flow. This takes mechanical work from the fluid and feeds it mechanically to the fan via a shaft. The generation of the drive power in the core flow and the conversion of this drive power into propulsion generation by the propulsion-generating fan in the bypass flow are functionally directly coupled according to the state of the art and integrated in close proximity to one another within a single engine housing.

In general, thrust as a force is physically defined as a change in momentum. In the case of thrust generation, in turn, this change in the impulse and thus the generation of a certain thrust can in principle take place on the one hand by accelerating a relatively small mass flow with a high change in speed. This possibility causes a high jet velocity in the outlet flow and thus a comparatively high jet noise. This type of propulsion generation is particularly suitable for achieving high flight speeds, for example in fighter jets.

On the other hand, the same thrust level can also be achieved by giving a relatively large mass flow only a relatively small increase in speed. This variant leads to a high propulsion efficiency and thus to an increased fuel efficiency at comparatively low flight speeds below Ma 0.9 and is therefore well suited for the economical propulsion of commercial aircraft.

In a modern turbofan jacketed-flow engine, the two mentioned possibilities of thrust generation occur in combination, with thrust generation generally dominated by a large mass flow, and even more so with increasing bypass flow ratio of the engine. In the core flow, only a small air mass flow is accelerated at a comparatively high speed and, exiting through the nozzle, primarily generates thrust by increasing the speed. In the bypass flow, on the other hand, the major part of the total air mass flow of the engine is accelerated by the fan at only low speed, ie here thrust is generated primarily by the movement of a large air mass flow with only a small additional speed distribution. The ratio of the air mass flow in the secondary flow and the air mass flow in the core flow is used as the secondary flow ratio or by-pass ratio, abbreviated to BPR, the engine. In the past, achieved improvements in specific fuel consumption are largely due to an increase in the bypass flow ratio.

In modern turbofan engines, the bypass flow generates the overwhelming majority of the thrust and at the same time, as a so-called "cold" jet, envelops the "hot" and "fast" jet of the core flow, which can be perceived as a noticeable reduction in noise when compared to single-flow engines. By increasing the bypass ratio, the noise has also been significantly reduced so far.

In the future, efforts are aimed at further increasing the bypass ratio of engines. As the influence of the secondary flow, which according to the design flows slowly, continues to increase, the necessary increase in the speed of the fluid, which must be given in an engine to generate a certain thrust level, also decreases overall, considered for the entire engine. To generate this, then overall low, increase in speed in the fluid, only a lower pressure gradient is now required in the fan that generates propulsion. The ratio of the pressure immediately downstream of the fan and that immediately upstream of the fan is referred to as the fan pressure ratio, or FPR for short. With a now lower fan pressure ratio, the fan or rotor noise emitted also decreases, which in modern bypass engines makes up a very significant part of the overall engine noise.

In summary, a certain thrust level is reached with an increased bypass flow ratio with a lower fan pressure ratio and an increased rotor area. As a result of this lower fan pressure ratio, the fluid experiences a lower overall increase in speed, which in turn increases the propulsion efficiency. With an increased propulsion efficiency, the specific fuel consumption also decreases proportionally. This results in lower mission fuel and increased fuel efficiency.

A very important point here is that with the same desired thrust level and approximately the same airspeed, this form of efficient thrust generation by means of a large mass flow requires an increase in the geometric thrust cross-sectional area swept by the fan blades in circulation. This thrust cross-sectional area of the propulsion-generating fan can also be called the rotor area and in many cases, for example with a propeller, also roughly corresponds to the inlet cross-sectional area through which a mass flow passes through the engine in order to participate in the generation of propulsion. In simplified terms, this is also easy to imagine. The mass flow that is to be increased here is defined as the mass of air that flows through a certain thrust cross-sectional area or rotor area in a certain time. The mass flow can also be defined as the product of the density of the air, the flow velocity and the thrust cross-sectional area through which this air flows to generate thrust. If the density of the fluid and the flight speed (which also roughly corresponds to the entry speed into the rotor surface) remain approximately the same in a first approach, the geometric rotor surface must be increased further in order to be able to achieve an unchanged thrust level. However, the rotor area can only be increased by increasing the diameter of the rotor generating the propulsion.

The following challenges arise with existing engines. As the fan diameter increases, the dimensions of the engine nacelle are inevitably larger with increasing bypass flow and increased propulsion efficiency. The larger engine nacelle, with its increased scavenged surface and interference effect, ensures greater drag. It should be noted that an engine nacelle is washed around on both sides by the fluid, creating resistance, both on its inside and on its outside. In this case, it is approximately the case that the resistance of the fan increases with the square of the fan diameter. This fact results in a critical catch size in a turbofan engine, or an associated critical bypass ratio from which the additional resistance of the engine due to its nacelle size consumes the improvement in specific fuel consumption achieved by increasing the bypass ratio. Starting from this limit, a further increase in the bypass flow ratio can then be expected to result in lower efficiency and a renewed increase in fuel consumption. Nowadays, bypass flow ratios of 12: 1 -13: 1 are already achieved in long-haul flight operations with conventional turbofan engines. A further significant increase in bypass flow and thus fuel efficiency is no longer possible with conventional engine architectures. There is therefore a need to create new engine architectures that allow a further increase in fuel efficiency.

Another problem with increasing bypass ratio is integrating engines with larger diameters on the aircraft. In the majority of today's commercial aircraft, the engine is configuratively mounted under the wings. For safety reasons, a certain ground clearance between the engine nacelle and the ground must be maintained. With increasing fuel efficiency, or with it increasing bypass flow ratio and fan diameter, a greater integration height on the aircraft between the ground and wing is necessary, taking into account a safety distance. The integration height available in practice on the aircraft, however, usually remains limited, for example by the structural wing connection and the wing box, which in the case of a flat passenger compartment floor must run underneath it. In the case of the Boeing 737, this is particularly evident for historical reasons. New constructive methods, such as the design of the wing as a "gull wing", help to a limited extent here, but the installation height generally remains geometrically limited. A further increase in the available installation height would also require an extension of the landing gear legs of the landing gear arrangement, which, however, would also lead to a high increase in weight, which would disadvantageously weaken the improved fuel consumption.

In addition, the circumferential speed reached at the fan blade tips increases proportionally to the growing fan diameter at a comparable angular rotation speed, given by the turbine running speed, whereby at higher critical values there is the risk that the efficiency of the propulsion conversion of the fan is reduced and that of the fan Noise generated by the blade tips increases unacceptably. In today's modern engines, these blade tips Ma numbers of up to 1.4 are in the supercritical range.

These challenges have led to the development of innovative engine architectures apart from conventional engines. With the Geared Turbofan GTF, the fan is decoupled from the speed of the drive and turbine shaft by means of a reduction gear. The lower rotation speed of the fan keeps the fan blade tip mach number low and increases its component efficiency. The engine can be designed for a lower fan pressure ratio, as a result of which the fluid in the bypass flow is less accelerated and flows more slowly. As a consequence, higher bypass ratios of up to 15: 1, higher propulsive efficiency and thus lower thrust-specific fuel consumption can be achieved with geared turbofan engines. Due to the lower speed distribution, a higher mass flow is necessary. Therefore, the thrust cross-sectional area, here the thrust-generating rotor area of the fan, must be increased. Due to the lower rotational angular speed of the fan, however, its diameter and thus its thrust surface can be increased without critical overspeed being reached at the blade tips. This enables the fan to operate with little noise, with the fan noise resulting approximately from the fourth power of the rotational speed of its blades. The improvement in specific fuel consumption is mainly achieved by improving the propulsion efficiency as a result of the increased bypass flow ratio. The disadvantage is that the rotor area of the propulsion-generating fan must nevertheless be enlarged and thus the engine nacelle is larger and more resistant to generation. The design as a geared turbofan brings with it an accepted additional engine weight of approximately 15-20%. With today's geared turbofans, savings in specific fuel consumption of 6-10% can be achieved, whereby the noise can be cumulatively reduced by 15-20 EPNdB, in particular due to the lower rotation speed of the fan.
If one wants to further increase the fuel efficiency of transport aircraft, it is necessary to use the so-called open rotor technology. As already mentioned before, the maximum fan diameter and therefore also the maximum propulsion efficiency is limited by the engine nacelle and the resulting drag, among other things.

Open rotor configurations have the potential to increase the propulsion efficiency even further and thus achieve significantly higher bypass flow ratios and significantly higher fuel efficiency. With the open rotor technology, the propulsion-generating rotor rotates in a free flow and is no longer surrounded by a cladding. This means that there is no need for the engine nacelle that generates resistance. In spite of this, open-rotor configurations according to the state of the art require an engine nacelle for the gas generator, which, however, is significantly smaller, has a diameter roughly the dimensions of the fairings of earlier single-flow engines and thus generates significantly less resistance than conventional engine nacelles and engine nacelles from Geared turbofans. In contrast to turbofan engines, open open rotor concepts do not contribute to an increase in drag due to the casing of the engine. In the past, at the time of the oil crises in the 1970s, open rotor technology was already under the name of that time Propfan technology tested on short-haul flying demonstrators in practice, with fuel savings of over 20% compared to engines at the time being proven.

The term propfan underlines that the rotor used for propulsion is the result of a combination of propeller and fan, whereby the high propulsion efficiency of a propeller is combined with the ability of the fan to work at high cruising speeds. As a result, it has been proven that high cruising speeds of up to Ma 0.86 can be flown with high fuel efficiency with open rotors. Rotors, the blades of which are aerodynamically designed similar to a prop fan, often have a small relative thickness and a continuous or progressive sweep towards the blade tip so that the critical Mach number is only reached when the inflow numbers at the blade tip are as high as possible. In addition, for the same reason, they can have a taper. In addition, progressive, transonic to supercritical profiles are often used in the profile section, also using the area rule. As a result of the measures mentioned, the critical Mach number at the blade tips is only reached at a relatively high speed of rotation of the blades, which means that at a relatively high cruising speed or blade flow, up to a Mach number of up to 0.86 at a cruising altitude of over 10,000 m can be flown. This airspeed is not possible to this extent with turboprop arrangements. Furthermore, the blades of propfan rotors are often characterized by a blade depth that is changed over the length of the blade and is increased overall. The common feature of the propfan or open rotor technology is that at least one open rotating rotor without casing is used for propulsion, which enables high bypass flow ratios of around or over 20: 1 and a significantly improved propulsion efficiency, often in conjunction with a enlarged rotor area and with the result a significantly improved fuel efficiency.

Particularly high efficiencies are achieved in particular by rotors rotating in opposite directions, so-called Counter Rotating Turbofans CRTF, which inter alia. By recovering the energy from the propeller spin, so-called swirl recovery, the efficiency of the rotors can be increased by a further 7-8% when cruising and by up to 12% when taking off. Thus, with previous designs, propulsion efficiencies of up to 92% in theory and up to 87% in practice in the installed state at the cruising speeds of Ma 0.78 desired for standard short-haul aircraft are achieved. The propulsion efficiency of today's engines of short-haul aircraft of the technology level of the year 2000 with only moderately high bypass flow ratios of 5-6 is only 0.70 - 0.72, here for the best case of cruise flight. Since the propulsion efficiency is directly included in the specific fuel consumption, previous open rotor configurations have the potential to improve fuel consumption significantly. In reality, it is assumed that a large part of this potential can be realized in practice and the specific fuel consumption can be improved by 15-20%, measured against the engines of the year 2000.

Two rotors rotating in opposite directions also have the potential to move a large mass flow through the engine with the same fan diameter and low speed. In addition, the lower fan pressure ratio means that fewer rotor blades are required in each stage, which can have a favorable effect on the engine mass and on noise development, in particular with regard to the noise caused by interference between the two rotors.

With the open rotor integration on the aircraft, three main challenges have not been resolved for a long time, two of which are causally related to one another. For the open rotor technology, an enlarged thrust cross-sectional area, i.e. rotor area, is necessary in order to be able to achieve significantly higher fuel efficiency with a high mass flow and low speed distribution through the rotor. According to the prior art, this enlarged rotor area is only possible by increasing the span, that is to say increasing the extent of the rotor blades. The angular speed of rotation of these blades is predetermined from the gas turbine process by the angular speed of the turbine stage which drives these rotor blades. With the large extension of the rotor blades and the resulting large rotor diameter, very high circumferential speeds result at the blade tips of the rotor. These lead to excessive speed-related noise emissions, which can reach a particular intensity in the radial direction within the rotor plane. Experience has shown that these are manifested in an disadvantageously increased external sound perception of the aircraft configuration due to the fan-like rotor, i.e. due to the so-called fan or rotor noise, especially during take-off and climb in the vicinity of the airport. Secondly, the supercritical and in some cases unsteady flows at the tips of the fan blades, which are connected to the noise, lead to vibrations in the rotor blades. Due to the large rotor diameter, this spreads favorably over the large blade extension, is sometimes intensified by vibration resonances due to the large blade length and finds its way through the fluid as well as noise as vibration and noise Via the structural connections of the engine in the airframe and passenger cabin. So far, the noise level has increased inside the cabin, sometimes also the level of vibration; outside the cabin, the noise level has so far increased significantly due to the rotor noise.

Thirdly, for safety reasons, present-day commercial aircraft have at least two independent engines, each with associated rotors, in accordance with the principle of redundancy. At the same time, it is a characteristic feature of the open rotor technology that the rotors can move freely in the medium and are therefore not surrounded by a casing or cladding in a radial outward direction. If a break or component failure occurs within the rotor, which must always be expected in the context of a sustainable safety assessment, parts of the rotor can be thrown out of the rotor by the effective centrifugal forces. Purely due to the architecture, such a case of damage always results in a so-called "Uncontained Engine Failure".

Due to the higher circumferential speed of the rotors compared to turbofan engines with open rotor engines, their comparatively larger diameter and the higher weight of the rotating components, those fractions (debris) that leave the rotor in the event of damage have increased kinetic energy is potentially suitable for breaking through components of the aircraft that are essential for safe flight, such as the printed passenger cabin, fuel and flight control systems, wings, tanks and control surfaces, etc. and thereby damaging them in a safety-relevant manner. This means that these circumstances must be taken into account when positioning these drive systems.

With regard to the drive system, it is also particularly relevant that parts of a rotor that escape in the event of damage cannot hit the rotor of a further, independent engine and also disable it, which would remove the redundancy principle of the aircraft. Previously presented aircraft configurations for open rotor drive integration often do not take sufficient account of this safety-relevant fact. Placing open-rotor engines directly next to one another, as in the case of turbofan engines, as shown in configuration studies according to the state of the art, harbors the risk that, in the "rotor burst trap" of an engine, both engines will be damaged by fragments of the rotors and thus the Drive redundancy of an aircraft is overridden.

According to the state of the art, the arrangement of open-rotor engines on the passenger aircraft has therefore been unsatisfactory overall for a long time, and the arrangement of the engines on the aircraft has hitherto been severely limited. In practice, with twin-engine transport aircraft, apart from individual research aircraft, only an unfavorable engine arrangement in the rear area of the unprinted tail is possible for safety reasons, with the rudder unit being placed between the two rotors for protection. In this way, in the event of damage, it was intended to prevent the second rotor from being damaged after a break in one rotor, which would otherwise have rendered an engine redundancy concept ineffective. It is controversial in technology whether the rudder unit could even be structurally designed in such a way that it could ensure effective shielding of the rotors in practice in the event of a fault. If this were to succeed in practice, the rudder unit would be very heavy for structural reasons. Furthermore, in the event of a rotor burst, with this arrangement, components of the aircraft that are necessary for safe flight can still be damaged in a safety-relevant manner, since, for example, even if the swept tail control surfaces are arranged outside the rotor plane, control and hydraulic lines move the rotor planes in the longitudinal direction have to cross.

For the aircraft, however, the rear arrangement of the engines shown so far is very unsatisfactory overall because it dictates not only the position of the wing arrangement, but also essentially the entire aircraft configuration via the engine position. Due to the rear engine position, the influence on the center of gravity is great, the loading of the aircraft with payload and fuel can therefore not be carried out as flexibly as it would be desirable, which reduces the operational flexibility of the aircraft. The tail lever arms are small, which is why the tail surfaces must be made large, which increases the resistance and weight. In addition, due to the position of the tail engine, the moments of inertia around the transverse axis of the aircraft are large, which further increases the surface area of the tail units. In addition, a comparatively heavy T-tail unit must be installed, which is even heavier if this is to enable structural shielding between the two open rotors, or if a trimmable horizontal stabilizer is to be installed, which is standard in commercial aircraft. The trim resistance can also be increased when cruising, due to the configuration of the center of gravity and tail lever arms. For safety reasons, the open rotors must be on one Longitudinal section of the aircraft are positioned along the longitudinal axis of the fuselage, where there is no longer a pressurized cabin, since otherwise the pressurized cabin could break through and be damaged by fragments in the event of a break. In addition, the open rotors with enlarged rotor surface with long, heavy and resistance-generating pylons must be attached to the tail of the aircraft. Although the rotors run open in the fluid, a smaller engine nacelle is still necessary, which further increases the area of the aircraft that is washed and thus contributes to the aircraft's zero and interference resistance. The problem of the center of gravity is exacerbated by the additional local weight in the rear area.

The main problem that has not yet been solved consists in finding a safe form of arrangement for open rotors for the aircraft in such a way that essential components of the aircraft cannot be damaged for safe flight. In addition, it has not yet been resolved to find a form of arrangement for open rotors in such a way that, apart from the performance improved by the engine, it also maintains or improves the performance and operational flexibility of the core aircraft.

According to the prior art, it is known to arrange propellers surrounding the fuselage of an aircraft.

To CH 235 699 A aircraft are known in which a propeller is built into the fuselage and several propellers rotating in opposite directions to each other can also be used on the fuselage. The propellers are arranged in the direction of flight behind the wing arrangement, so that the arrangement of weapons on the aircraft is easier and so that the “propeller circle” does not have to be shot through. The aircraft does not have a pressurized cabin inside the fuselage. The motor device is mounted inside an aircraft fuselage and, via a pinion, which for the most part rotates within the fuselage boundary, drives a larger diameter circumferential ring supported by a bearing ring. The drive is achieved here via an internally toothed gear attached to this circumferential ring. Since the one bearing ring with a smaller diameter is rigidly connected to the fuselage and this is closed over the full angular circumference, i.e. not interrupted, in this arrangement the motor pinion must break through the fuselage outer skin via an opening along the longitudinal axis of the fuselage in front of or behind the bearing ring in order to be able to use to come into contact with the internally toothed circumferential ring of larger diameter. Another disadvantage is that the engine is installed inside the fuselage arrangement, with no separate air supply and exhaust gas discharge, which makes the long-term operation of air-breathing engines difficult and also takes up space within the fuselage arrangement, which is no longer available for any transport tasks .

An additional important disadvantage is that the propeller is arranged at the tail area of the aircraft. This means that turbulences in the wing get into the propeller and, as experience shows, when interacting with it, e.g. as is known from the Piaggio Avanti, through increased noise emitted by the propeller. Experience has shown that the aircraft will therefore have an increased noise intensity in terms of external perception during operation.

Another disadvantage is that the tail unit at the stern has to be made unsuitably small so that the propeller can be dismantled towards the rear. This also has the consequence that the tail units are no longer irradiated by the propeller jet. In general, it is a disadvantage of rotary propellers in the rear position that they are difficult to remove from the aircraft due to the tail units at the rear behind the propeller and due to the wing in the front and therefore have to be dismantled.
DE 18 84 174 U refers to a powered wing aircraft, which applies both the lift to carry the aircraft and the propulsion to propel the aircraft via rotating rotors of large radial extension, which can be pivoted kinematically with parts of the fuselage and surround parts of the fuselage. The rotors are driven by assigned blade tip drives. The aircraft can take off and land vertically. However, the disadvantage of this arrangement is the high complexity of this configuration and sensitivity in the event of a failure of a subsystem, for example the swivel mechanism of at least one rotor. The rotors are designed to be quite large in their radial length, so that a safe landing is no longer possible if a pivoting mechanism fails. In addition, with this aircraft, due to the blade tip drive and the ability to take off and land almost vertically thanks to rotating rotors, a high absolute and performance-specific energy consumption as well as a high level of noise intensity during operation can be expected. In the event of a rotor breaking, fragments of high energy escaping can damage essential components of the aircraft for flight safety, such as the landing gear arrangement and tail units. In addition, if a rotor system fails, the aircraft is no longer airworthy.

The pamphlets too GB 2,468,917 A ; WO 2012/114 047 A1, DE 44 26 403 C1 and EP 1 428 752 A1 show different ways of attaching rotors. However, it is not stated that a pressurized cabin is used on the aircraft or that the designs and arrangements of rotors are technically associated with the pressurized cabin.

Object of the invention

The object of this invention is to find a safe arrangement, embodiment and drive form for bladed rotors for an aircraft configuration, which is also suitable in the sense of the open rotor concept for drive rotors freely rotating on the outside and for drive rotors with high kinetic energy, and which is characterized by that the fragments accelerated by centrifugal force in the event of a rotor breakage are less likely to hit and damage other structural components of the aircraft that are important for the safe execution of the flight, in particular also the pressurized cabin.

Another object of this invention is to find solutions for propulsion rotors and propulsion systems on aircraft that enable a high bypass ratio up to the high cruising speeds of turbo-powered commercial aircraft that are customary today, as well as reducing the noise emitted during propulsion and sustainably reducing the fuel efficiency of the propelled aircraft to increase.

Inventive solution

• - einer zumindest abschnittsweisen zylindrisch gestreckten Rumpfanordnung F mit einer räumlichen Haupterstreckung in Richtung einer Rumpflängsachse FA, diese Rumpfanordnung F zur bedarfsweisen Unterbringung von Nutzlast, wobei die Rumpfanordnung F zumindest abschnittsweise in der Funktion einer Druckkabine P mit einem Innendruck bedruckbar ist,
• - einer fest an die Rumpfanordnung F des Flugzeugs angebundenen Tragflächenanordnung W, zur Generierung eines das Flugzeug tragenden Auftriebs, wobei die im Horizontalflug erzeugte Kraftkomponente des Auftriebs senkrecht zur Rumpflängsachse FA gegenüber der Komponente längs zur Rumpflängsachse FA deutlich überwiegt,
• - mit einer bedarfsweise einziehbaren und wieder ausfahrbaren Dreipunktfahrwerksanordnung GA mit mehreren drehbar gelagerten Rädern WE zur statisch bestimmten Abstützung des Flugzeuges entgegen der Bodenebene BO im bodenbündigen Betrieb, bestehend aus zwei strukturell an die Tragflächenanordnung W angebundenen Hauptfahrwerksbeinen GAM und einem an der Rumpfanordnung F strukturell angebunden Bugfahrwerksbein GAA, wobei das Bugfahrwerksbein GAA entlang der Rumpflängsachse FA in Flugrichtung FR vor den Hauptfahrwerksbeinen GAM angeordnet ist,
• - einer Motorenanordnung E, bestehend aus wenigstens einem Motor, zur Generierung einer dem Flugzeug zur Verfügung stehenden Antriebsleistung,
• - wenigstens einem mit mehreren Rotorschaufeln B beschaufelten Rotor R zur zumindest teilweisen Umwandlung von Antriebsleistung der Motorenanordnung E in eine das Flugzeug treibende Vortriebsleistung durch Rotation des beschaufelten Rotors R um eine geometrische Rotationsachse RA, wobei sich bei Rotation der Rotorschaufeln B des beschaufelten Rotors R im Umlauf geometrisch eine von den Rotorschaufeln B überstrichene geometrische Rotorfläche RAE ergibt, dieser beschaufelte Rotor R in seiner radialen räumlichen Erstreckung einen kleinsten Innendurchmesser ID und einen größten Außendurchmesser OD aufweisend, wobei dieser beschaufelte Rotor R mit im Einstellwinkel verstellbar ausgeführten Schaufeln B ausgerüstet ist, dieser beschaufelte Rotor R ebenfalls eine räumliche Längserstreckung entlang seiner geometrischen Rotationsachse RA aufweisend, wobei innerhalb dieses Bereiches seiner Längserstreckung Rotorebenen RP senkrecht zu seiner geometrischen Rotationsachse RA als geometrisches Hilfsmittel kennzeichenbar sind, um einen möglichen Auswirkungsbereich dieses Rotors R zu erfassen,
• - wenigstens einem Getriebe G zur Übertragung der Antriebsleistung wenigstens eines Motors der Motorenanordnung E auf wenigstens einen beschaufelten Rotor R,
• - wenigstens einer Lagerandordnung S,

dadurch gekennzeichnet, dass wenigstens ein beschaufelter Rotor R mit mindestens einer Lageranordnung S, drehbar die Rumpfanordnung F umgebend gelagert ist sowie, dass mindestens dieser eine beschaufelte Rotor R ungefähr in Richtung der Rumpflängsachse FA durch wenigstens eine Lageranordnung S zur Rumpfanordnung F axial fixiert ist, und dass mindestens dieser eine beschaufelte Rotor R zugleich in radialer Richtung außen dabei denjenigen Abschnitt einer zumindest abschnittsweisen zylindrisch gestreckten Rumpfanordnung F eines Flugzeuges ringförmig umgibt, welcher, in einer der Rotorebenen RP dieses Rotors R betrachtet, zumindest anteilig in der Funktion einer Druckkabine P bedruckbar ist und, dass zumindest in einer der Rotorebenen RP des beschaufelten Rotors R keine weiteren für den unmittelbar sicheren Betrieb des Flugzeuges wesentlichen Komponenten des Flugzeuges angeordnet sind, und dass mindestens dieser eine beschaufelte Rotor R entlang der Rumpflängsachse FA, in Flugrichtung FR gesehen, vor der Tragflügelanordnung W angeordnet ist und zugleich die geometrische Rotationsachse RA mindestens eines beschaufelten Rotors R, entlang einer Richtung der Flugzeughochachse VA, ausgehend von der Bodenebene BO aus gesehen, in einer Ebene betrachtet, die unmittelbar stromaufwärts der Tragflächenanordnung W und senkrecht zur Rotationsachse RA dieses beschaufelten Rotors R liegt, so angeordnet ist, dass sie sich oberhalb des sich in dieser Ebene ergebenden minimalen Vertikalniveaus MV der oberen Oberfläche der Tragflügelanordnung W, befindet so, dass in einer Rotorebene RP dieses einen beschaufelten Rotors R der überwiegende Flächenanteil, der im Umlauf von den Rotorschaufeln B des beschaufelten Rotors R überstrichenen geometrischen Rotorfläche RAE, von der Bodenebene BO aus betrachtet, oberhalb der Tragflächenanordnung W zum Liegen kommt so, dass die auf die, entgegen der Flugrichtung FR nachgeordnete Tragflächenanordnung W bewirkte Staudruckerhöhung des Propellerstrahls mindestens dieses einen beschaufelten Rotors R überwiegend oberhalb der Tragflächenanordnung W erfolgt und, dass mindestens dieser eine Rotor R in Richtung der Rumpflängsachse FA an der Rumpfanordnung F zwischen dem Bugfahrwerksbein GAA und den Hauptfahrwerksbeinen der Hauptfahrwerksanordnung GAM so angeordnet ist, dass, in jeder der Rotorebenen RP dieses einen beschaufelten Rotors R betrachtet, der größte Außendurchmesser OD dieses einen Rotors R kleiner ist als der, mit durch die Fahrwerksanordnung GA vorgegebene, doppelte minimale Abstand der Rotationsachse RA dieses beschaufelten Rotors R von der Bodenebene BO, und, dass mindestens ein Motor der Motoreneinrichtung E in einem Bereich des Flugzeuges außerhalb der Druckkabine P, aber zumindest zum Teil innerhalb eines Raums versenkt angeordnet ist, der nach außen zum Fluid durch die von der Außenhaut der Tragflächenanordnung W und die von der Außenhaut der Rumpanordnung F gebildeten Außenkontur des Flugzeuges abgegrenzt wird, und so in diesem Raum eingebettet angeordnet ist, dass ein Bauteil, das in der Funktion einer Motorenhalterung strukturelle Kräfte von mindestens einem Motor der Motorenanordnung E in Richtung der Struktur des Flugzeug abführt, im Wesentlichen nicht von Außen ersichtlich ist und, dass diesem Bauteil in seiner Funktion der Motorenhalterung außerhalb der durch die von der Außenhaut der Tragflächenanordnung W und von der Außenhaut der Rumpfanordnung F gebildeten Außenkontur im Wesentlichen keine eigene vom Fluid bespülte, widerstandswirksame Oberfläche zuordbar ist, und, dass mindestens dieser eine beschaufelte Rotor R mit wenigstens einem Motor der Motorenanordnung E, der außerhalb der Druckkabine P angeordnet ist, über Getriebe F, kinematisch gekoppelt ist, und durch diesen in Rotation versetzt werden kann, um eine, das Flugzeug treibende Vortriebskraft zu erzeugen, dessen Kraftkomponente in Richtung der Rumpflängsachse FA gegenüber der Kraftkomponente senkrecht zur Rumpflängsachse FA deutlich überwiegt.The task is solved by an airplane with:

• - An at least partially cylindrically elongated fuselage arrangement F. with a spatial main extension in the direction of a longitudinal axis of the trunk FA , this hull arrangement F. for the storage of payload as required, the fuselage arrangement F. at least in sections in the function of a pressurized cabin P. can be printed with an internal print,
• - one fixed to the hull assembly F. of the aircraft attached wing assembly W. , for generating a lift carrying the aircraft, the force component of the lift generated in horizontal flight being perpendicular to the longitudinal axis of the fuselage FA opposite the component along the longitudinal axis of the fuselage FA clearly predominates,
• - With a three-point landing gear arrangement that can be retracted and extended again if necessary GA with several rotatable wheels WE for statically determined support of the aircraft against the ground plane BO in ground-level operation, consisting of two structurally attached to the wing assembly W. attached main landing gear GAM and one on the fuselage assembly F. structurally attached nose gear leg ATM , being the nose gear leg ATM along the longitudinal axis of the trunk FA in the direction of flight FR in front of the main landing gear GAM is arranged
• - a motor arrangement E. , consisting of at least one motor, to generate a drive power available to the aircraft,
• - At least one with several rotor blades B. Bladed rotor R for at least partial conversion of drive power of the motor arrangement E. into a propulsive power that drives the aircraft by rotating the bladed rotor R around a geometric axis of rotation RA , with rotation of the rotor blades B. of the bladed rotor R in rotation geometrically one of the rotor blades B. swept geometric rotor area RAE results, this bladed rotor R has a smallest inner diameter in its radial spatial extension ID and a largest outside diameter OD having, this bladed rotor R with adjustable blades designed in the setting angle B. is equipped, this bladed rotor R also has a spatial longitudinal extent along its geometric axis of rotation RA having, within this region of its longitudinal extent rotor planes RP perpendicular to its geometric axis of rotation RA can be identified as a geometric aid in order to capture a possible area of impact of this rotor R,
• - at least one gear G for transmitting the drive power of at least one motor of the motor arrangement E. on at least one bladed rotor R,
• - at least one storage arrangement S. ,

characterized in that at least one bladed rotor R with at least one bearing arrangement S. , rotatable the fuselage assembly F. is mounted surrounding and that at least this one bladed rotor R approximately in the direction of the longitudinal axis of the fuselage FA by at least one bearing arrangement S. to Hull arrangement F. is axially fixed, and that at least this one bladed rotor R at the same time in the radial direction outside that section of an at least partially cylindrically elongated fuselage arrangement F. of an aircraft, which, in one of the rotor planes RP this rotor R considered, at least partially in the function of a pressurized cabin P. is printable and that at least in one of the rotor planes RP of the bladed rotor R no further components of the aircraft essential for the immediate safe operation of the aircraft are arranged, and that at least this one bladed rotor R along the longitudinal axis of the fuselage FA , in the direction of flight FR seen in front of the wing assembly W. is arranged and at the same time the geometric axis of rotation RA at least one bladed rotor R, along a direction of the aircraft vertical axis VA , starting from the ground plane BO when viewed in a plane immediately upstream of the wing assembly W. and perpendicular to the axis of rotation RA this bladed rotor R is arranged so that it is above the minimum vertical level resulting in this plane MV the top surface of the wing assembly W. , located so that in a rotor plane RP This one bladed rotor R is the predominant area that is in circulation by the rotor blades B. of the bladed rotor R swept geometric rotor surface RAE , from the ground level BO viewed from above the wing assembly W. comes to rest in such a way that it faces, against the direction of flight FR downstream wing assembly W. caused a dynamic pressure increase of the propeller jet at least this one bladed rotor R predominantly above the wing arrangement W. takes place and that at least this one rotor R in the direction of the longitudinal axis of the fuselage FA on the fuselage assembly F. between the nose gear leg ATM and the main landing gear legs of the main landing gear assembly GAM is arranged so that, in each of the rotor planes RP this considering a bladed rotor R, the largest outside diameter OD this one rotor R is smaller than the one with by the undercarriage arrangement GA specified, double the minimum distance of the axis of rotation RA this bladed rotor R from the ground plane BO , and that at least one motor of the motor device E. in an area of the aircraft outside the pressurized cabin P. , but is arranged sunk at least in part within a space leading to the outside to the fluid through the from the outer skin of the wing assembly W. and that of the outer skin of the hull assembly F. formed outer contour of the aircraft is delimited, and is so embedded in this space that a component that acts as an engine mount structural forces of at least one engine of the engine assembly E. in the direction of the structure of the aircraft, is essentially not visible from the outside and that this component in its function of the engine mount outside of the by the outer skin of the wing assembly W. and from the skin of the fuselage assembly F. formed outer contour essentially no separate, drag-effective surface flushed by the fluid can be assigned, and that at least this one bladed rotor R with at least one motor of the motor arrangement E. outside the pressurized cabin P. is arranged via gear F. , is kinematically coupled, and can be set in rotation by this in order to generate a propulsive force driving the aircraft, the force component of which is in the direction of the longitudinal axis of the fuselage FA compared to the force component perpendicular to the longitudinal axis of the fuselage FA clearly predominates.

The special mode of operation of the invention arises from a special combination of individual or several inventive aspects.

According to an important aspect of the invention, at least one bladed rotor R is thus a fuselage arrangement that can be printed at least in sections F. , this ring-shaped surrounding rotatably arranged that the rotor R in one of its rotor planes RP , which in the area of its length along its rotor axis RA are arranged perpendicular to this, at least part of a pressure cabin P. that is inside the fuselage assembly F. lies, surrounds. According to the invention, this has the consequence that in the event of a rotor breakage during operation, fragments of the rotor R move away from the fuselage arrangement F. , but especially the pressurized cabin P. , remove by the effective centrifugal forces. As a result, according to the invention, these fragments can pass through the pressurized cabin P. and hull arrangement F. with their high kinetic energy, they no longer cause critical safety damage. This is both on the ground in the unprinted state of the pressure cabin P. the case, as well as in the printed state of the pressure cabin P. in cruise. According to the invention, the rotor surrounds in one of its rotor planes RP seen, that cross section of the fuselage assembly F. , which in this plane, at least in sections according to a pressure cabin P. is printable. This means that, for example, only one area in the fuselage is used as a pressurized cabin P. can be designed, for example, only a passenger cabin in the upper area of the fuselage cross-section, while there is an unprinted area for cargo, for systems, engines, landing gear or the like in the lower area. In a further important embodiment, this is the function of a pressurized cabin P. printable section now really printed with an overprint so that the print cabin P. inside the fuselage arrangement, an overpressure to an area outside the pressurized cabin P. has inside or outside the aircraft, for example to that area outside the fuselage arrangement F. , in which at least one rotor R rotates in the fluid.

The bladed rotor R is rotatable in relation to the fuselage arrangement F. by at least one bearing arrangement S. stored and approximately in the direction of the longitudinal axis of the fuselage FA by at least one bearing arrangement S. fixed arranged. Approximately to the direction of the longitudinal axis of the trunk FA in this context means that the bladed rotor R also has a normal drive camber (approximately up to a maximum of 12 °) in relation to the longitudinal axis of the fuselage FA may have in one plane or also spatially.

In addition, it is a further inventive feature that one or more rotors R so along the longitudinal axis of the fuselage FA on the fuselage assembly F. are arranged that in all of the rotor planes RP of a rotor that is perpendicular to the axis of rotation RP of the bladed rotor R and are arranged within its longitudinal extent, no further important safety-critical components necessary for the immediately safe flight of the aircraft are arranged. Thus, in the event of a bladed rotor R breaking, these cannot be damaged in a safety-critical manner.

The components that are important for the immediate safe execution of the flight can include, for example, the control surfaces, the spoilers, the high-lift system, the wing arrangement, fuel and flight control systems and important energy distribution systems with their lines and reservoirs, possibly also tail and landing gear arrangements and of course the pressure cabin, tanks and other engines count. In case of doubt, you can orientate yourself here on the approval regulations of the aircraft.

An alternative method here leads to a further embodiment of the invention. In order to find such a suitable arrangement position on the fuselage for open rotors, one could as a possibility, e.g. in the plane of a wing, a cone with the cone limitation, starting from the extension limits of the bladed rotor R with a certain defined and suitable opening cone, e.g. 2 x 15 °, which can also be referenced in more detail using suitable statistics and investigations, so that it appears statistically probable that the trajectory of fragments that leave the rotor R under the effect of centrifugal forces in a suitable operating state in the event of breakage is within exactly this cone lies. If this cone is now allowed to rotate around the axis of rotation of the rotor in space by the angular range of 360 °, a torus-like body of rotation of infinite radial external extension results. This is guaranteed by mentally moving the bladed rotor R along the fuselage F. Together with the stationary and directionally fixed virtual torus body, so that neither wings nor tail units nor other directly important parts of the aircraft are located within this virtual body, a safe position has been found for the rotor R, in its place the rotor on the fuselage arrangement F. , rotatably mounted, can be attached. In this way, several bladed rotors R can also be placed along the fuselage F. place safely.

For safety reasons, larger aircraft have several, often two, engines based on the principle of redundancy. Since fragments fly radially outwards faster in the event of a fault due to the large centrifugal forces with a suitable material and mass-appropriate design than they drift backwards in the air flow downstream, several are preferably independent, along the fuselage F. Bladed rotors R or drive systems staggered one behind the other are possible without these mutually influencing one another in the event of a fault. This solves a very fundamental reservation and disadvantage of open rotor drive systems and rotors, namely that of the safe arrangement and placement on the aircraft, which has hitherto prevented practical use.

In a further possible and alternative embodiment, the rotor R could be designed to be encased in its circumference radially on the outside, at least over a certain angular range of the rotor R, in the circumferential direction by a cladding. This cladding can also be designed to be retractable, for example so that it can be retracted or folded in towards the rear in the direction of flow so that it fits aerodynamically as well as possible into a surrounding fuselage contour.

The fairing can advantageously increase the engine thrust given off at low operating speeds, especially during take-off. In addition, this cladding can help to reduce the noise radiated downwards and to the side during take-off and landing by means of shielding. At cruising speed, on the other hand, the power efficiency of the rotor is higher when the casing is retracted and sunk and no longer surrounds the bladed rotor R at least partially in the circumferential direction.

Alternatively, however, the casing could also be attached directly to the outside of the rotor R and thus move with it while rotating.

According to a particular characteristic of the invention, at least one bladed rotor R is designed so that it rotates freely in the fluid without a jacket. As a result, it achieves a very high level of efficiency in the propulsion conversion.

According to the invention, the bladed rotor R is designed overall so that its rotor blades B. are adjustable in the angle of incidence to the flow. This allows the high efficiency of the bladed rotor R to be used over a wide speed range of the aircraft. The change in the setting angle on the rotor blades is carried out by conventional mechanical devices in conjunction with actuators. It can also be any rotor blade B. a separate actuator must be assigned.

In a further very preferred development of the invention, two bladed rotors R are installed, which rotate in opposite directions to one another. Together they form a rotor system. In this way, the rotor efficiency of the rotor system is improved by 7-8% when cruising and by up to 12% when taking off compared to a single rotor. This contributes to a further advantageously improved efficiency of the aircraft. In addition, due to the rotation of the two rotors R, the aircraft does not experience a reaction torque on one side, since the reaction torques of the two rotors compensate each other.

In a further embodiment, in addition to these two bladed rotors R, one or more further pairs of rotors R rotating in opposite directions can be used at any fuselage position, the fuselage arrangement F. surrounding, be arranged. In addition, due to the rotation of a rotor R, the aircraft does not experience a reaction torque on one side. This is compensated by the operation of two bladed rotors R with different directions of rotation relative to one another.

Another additional feature essential to the invention is given by the fact that at least one bladed rotor R is attached to the fuselage arrangement F. not only this is arranged surrounding it, but in a certain way compared to a 3-point chassis arrangement GA with a nose landing gear ATM , which with the statically determined support of the aircraft against the ground level BO serves in operation flush with the ground, is positioned so that at least one bladed rotor R on the fuselage arrangement along the longitudinal axis of the fuselage FA between a nose landing gear structurally attached to the fuselage assembly ATM and a main landing gear assembly GAM is positioned, this being attached to the airfoil assembly W. is structurally connected and includes two main landing gear legs. Since the rotation during take-off and the derotation during landing are essentially centered around the wheels of the main landing gear assembly GAM takes place, there is therefore no risk according to the invention during these critical phases that a bladed rotor R on the ground level BO touches down or its ground clearance is impermissibly reduced during the rotation or derotation of the aircraft. This reduction would also not be advantageous in view of any damage caused by foreign bodies (FOD) which could be whirled up during the take-off run and could get into the bladed rotor R. A set-up of a bladed rotor R on the ground level BO during the rotation could sometimes have fatal consequences. On the other hand, according to a consequence of the invention, it is even the case that the available ground clearance of the bladed rotor R relative to the ground plane BO increased during these critical operating phases. In addition, there are the landing gear legs of the main landing gear arrangement GAM in the opposite direction of flight RFD , thus approximately in the direction of flow, behind at least one rotor R so that the inflow of the rotor in the extended state of the landing gear arrangement cannot be disturbed by the main landing gear legs. Thus, starting from this, no turbulence occurs in the rotor plane, which experience has shown could interact with the bladed rotor R in such a way that the rotor noise would be increased. This promotes low-noise operation, especially during the sensitive phase of take-off and landing. Turbulence can still occur, starting from the extended nose landing gear arrangement ATM reach the bladed rotors R, but the nose landing gear leg is aerodynamically easier to clad in such a way that this turbulence is minimized, for example by a light cladding that surrounds the nose landing gear leg, or also by a wheel cladding similar to a so-called wheel "slipper". This cladding can also be combined with a deflector known from other aircraft, for example the Concorde main landing gear, to prevent foreign objects from passing through the nose landing gear arrangement ATM can be whirled up. In addition, the nose landing gear leg is preferably further away from the rotors R and the aerodynamic caster area can also be located outside the area of influence of the rotor surface with an increased angle of attack during take-off and landing.

In a further possible embodiment, it would also be conceivable to use the nose landing gear arrangement ATM to be arranged downstream of the bladed rotor R, that is to say following it. According to a further advantageous aspect of the invention, in the wake of a landing gear leg, in particular the nose landing gear leg, in the direct line of sight between the landing gear leg and the bladed rotor R, a Lowerable or extendable shielding protective devices are provided, which make it difficult for foreign bodies whirled up by the chassis to subsequently get into the bladed rotor R and damage it. The extension can take place mechanically, electrically, hydraulically or with the help of the dynamic pressure or as a function of the weight. Alternatively, a kind of protective plate could also be attached to marked wheels of the landing gear or on the landing gear flaps, which prevents foreign bodies from being whirled up by the landing gear and getting into the rotor.

Another additional inventive feature is that at least one bladed rotor R between a nose gear arrangement ATM and a main landing gear assembly GAM in the direction of the longitudinal axis of the fuselage FA on the fuselage assembly F. is arranged so that in each of the rotor planes RP this considering a bladed rotor R, the largest outside diameter OD this one bladed rotor R is smaller than the one with the undercarriage arrangement GA specified, double the minimum distance of the axis of rotation RA this bladed rotor R from the ground plane BO . In this way, the bladed rotor R can safely rotate in all of the usual phases of operation flush with the ground, without being flush with the ground BO to get in touch. It can also be advantageous to use the largest outside diameter OD of a bladed rotor should be selected so that the bladed rotor R also has a safety distance to the ground level BO having. This can be sensible from a safety point of view, for example around 60 cm, or it can also be based on the geometry of the wheels of the nose landing gear arrangement ATM reference. The distance between the bladed rotor R in one of the rotor planes RP of the bladed rotor R is due to the geometry of the chassis arrangement GA , the location of the rotation thing RA of the bladed rotor R and the geometry of the fuselage arrangement F. can be influenced in their cross-section.

According to the prior art, it is the case that the bladed rotor R is arranged downstream of the wing assembly, surrounding the fuselage.

A very essential further aspect of the invention is that at least one rotor R, preferably also two of them, in relation to the wing arrangement W. in the direction of flight FR is positioned in front of this.

This means that the aerodynamic flow onto the rotor R cannot pass through the wing arrangement W. be disturbed. Thus, in contrast to the prior art, there is no turbulence in the lift-generating wing W. into the bladed rotor R, which experience has shown would greatly increase the noise emitted by the bladed rotor R. In addition, bladed rotors R arranged in front of the wing experience a largely homogeneous speed distribution in their flow over their rotor surface. This would not be the case if a bladed rotor R according to the invention were positioned in accordance with the prior art behind the airfoil, since the airfoil arrangement W. on the top and bottom to generate lift in the fluid induces a different speed distribution and the bladed rotor R, this wing arrangement W. downstream, due to the extent of its geometric rotor surface RAE , includes both the wake of the wing on the upper side and the aerodynamic wake of the wing underside, both of which have strong pressure and speed differences to one another, which is proven by known tip vortices that emanate from the wing tips as wake vortices.

By placing the bladed rotors R in front of the wing assembly W. a further low-noise operation of the aircraft succeeds according to the invention. Are at the same time, a very preferred embodiment of the invention, two bladed rotors R with opposite directions of rotation in front of the wing arrangement W. arranged, the flow on the wing arrangement succeeds W. In contrast to the prior art, it is also advantageous that it is largely twist-free.

Another important characteristic of the invention is that the geometric axis of rotation RA at least one bladed rotor R, viewed in a plane immediately upstream of the airfoil assembly W. and perpendicular to the axis of rotation RA this bladed rotor R lies, and along the direction of the aircraft vertical axis VA from the ground level BO seen from above the resulting minimum vertical level in this plane of the upper surface of the wing assembly W. , is arranged so that in a rotor plane RP This one bladed rotor R is the predominant area that is in circulation by the rotor blades B. of the bladed rotor R swept geometric rotor surface RAE , from the ground level BO viewed from above the wing assembly W. comes to rest in such a way that those facing in the opposite direction of flight RFD downstream wing assembly W. caused ram pressure increase of the propeller jet at least this bladed rotor R predominantly above the wing arrangement W. he follows.

Because the bladed rotor R is in front of the airfoil arrangement in the direction of flow W. is attached, the generation of propulsion succeeds according to the invention not only with little noise, but the propeller jet necessary for propulsion also contributes at the same time to an increased generation of lift on the wing arrangement W. at. This further synergy effect according to the invention further advantageously increases the efficiency of the aircraft. The background to this is that a rigid wing arrangement causes its lift by inducing a speed difference in the fluid between the upper side of the wing and the lower side of the wing, which causes a pressure difference between the upper and lower side. According to a simplified rule of thumb, 2/3 of the lift generated on a conventional wing is due to a negative pressure on the upper side and 1/3 to an overpressure on the underside of the wing. This rule of thumb does not always apply, as the exact geometry of the wing is also decisive, but the tendency is correct that the overspeeds generated by the wing on the upper side of the wing in relation to the free flow are greater than the change in speed and thus also the change in pressure on the underside of the wing.

According to the invention, the power introduced by the bladed rotor R and thus the induced increase in speed in the fluid, that is to say the increase in dynamic pressure, is introduced into the fluid predominantly above the wing, which increases the lift. According to the invention, this is achieved in that the geometric axis of rotation RA at least one bladed rotor R, viewed in a plane immediately upstream of the airfoil assembly W. and perpendicular to the axis of rotation RA this bladed rotor R lies, and along the direction of the aircraft vertical axis VA from the ground level BO seen from above the resulting minimum vertical level in this plane of the upper surface of the wing assembly W. , is arranged so that in a rotor plane RP This one bladed rotor R is the predominant area that is in circulation by the rotor blades B. of the bladed rotor R swept geometric rotor surface RAE , from the ground level BO viewed from above the wing assembly W. comes to rest in such a way that those facing in the opposite direction of flight RFD downstream wing assembly W. caused the ram pressure increase of the propeller jet at least this one bladed rotor R predominantly above the wing arrangement W. he follows.

If a bladed rotor R is designed as efficiently as possible, then this has its geometric rotor surface in its outflow RAE seen, over a speed field that is as uniform and homogeneous as possible. The bladed rotor R can therefore have a very similar outflow between the top and bottom of the rotor surface RAE can be assumed. This also takes place from the rotor R in front of the wing arrangement W. Power induced in the fluid, the predominant area of the geometrical rotor surface RAE accordingly, also predominantly on the top of the wing assembly W. which increases the lift of the aircraft in addition to the generated propulsion. This is advantageous when cruising because the required angle of attack of the aircraft can be selected to be smaller and thus less resistance for the aircraft arises, which increases efficiency. During take-off and climb as well as go-around, i.e. in the flight phases in which a lot of lift is required, more lift can be generated by the aircraft at a certain angle of attack, which also corresponds to an increase in performance.

A very important advantage according to the invention also results from the fact that the rotor or rotors R are in the direction of flight FR the wing assembly W. are connected upstream and likewise the tail assembly LW of the wing assembly W. is downstream in the direction of flow. This allows the rotor R, aided by the retractable landing gear of the landing gear arrangement GA , for example, for maintenance, repair or replacement, can be pulled forward over the fuselage, even in one piece. It no longer has to be dismantled for this.

In a further preferred embodiment, the aircraft has a tail unit arrangement, this consisting of a horizontal tail unit arrangement and a vertical tail unit arrangement. This tail arrangement is in the direction of the longitudinal axis of the fuselage FA , in the direction of flight FR seen, in front of and in the direction of flow behind the wing assembly W. so installed according to the invention that both the horizontal stabilizer arrangement and the vertical stabilizer arrangement are arranged at least in sections in the direct wake of the propeller jet generated by at least one bladed rotor R. The tail units are at least partially exposed to an increased dynamic pressure while lying in the propeller jet of a rotor and are therefore more effective. Under certain circumstances, due to their improved effectiveness, they can also be designed with a smaller tail surface, i.e. with lower resistance overall.

This is particularly advantageous during take-off in the rotating phase, because here the tailplane is still flown against at a comparatively low speed. In the case of cross winds, on the other hand, it is advantageous if the rudder assembly also works more effectively.

In general, the direct wake of a propeller jet generated by a bladed rotor R can be determined by making straight lines parallel to the direction of flight FR each of the radial extension limits of a bladed rotor R are drawn downstream. Alternatively, these straight lines can also be based on the direction of the undisturbed aerodynamic inflow in front of the aircraft.

In general, it is a further advantage of the invention that, if several rotors are attached to the aircraft according to the invention, there is no moment about the vertical axis during take-off run on the ground in the event of a failure of one of these or of a motor assigned to them VA of the aircraft occurs. In the event of a motor or rotor failure, the tail unit arrangement also remains energized and, due to the central arrangement of the rotors R, does not have to intervene significantly in a compensatory manner with the help of control movements of the rudders, which means that the surface of the tail units of the tail unit arrangement LW is smaller, i.e. also has less resistance could be interpreted.

According to a further development of the invention, individual rotor blades B. at least one bladed rotor R can be adjusted independently of one another in their setting angle to the air flow, and also vary over the circumferential position of the rotor, so that a vectorization of the thrust results, which is used for control, trimming, drag saving and noise reduction of the aircraft can be. However, as described above, the special inventive aspect of the additional increase in lift by at least one bladed rotor R can be additionally reinforced in this way. The pitch angle of each rotor blade B. is independent of the setting angles of other blades B. can be changed, preferably by assigned actuators, according to an individual blade control concept.

According to an advantageous development of the invention, there is the motor device E. of several motors, with each bladed rotor R for power supply at least one specific motor of the motor device E. assigned. In this way, the redundancy requirements can also be met, and thus, for example, a bladed rotor R from a motor of the motor arrangement E. and a further bladed rotor R, which can also rotate in opposite directions, from a second motor of the motor arrangement E. are driven.

In another embodiment, there is the engine assembly E. from several motors, one being the rotor plane RP perpendicular to its axis of rotation RA a bladed rotor R considered, these motors are arranged so that at least one motor in an area between the nose landing gear leg ATM and the right main landing gear leg of the main landing gear assembly GAM and at least one second motor in an area between the nose gear leg ATM and the left one of the main landing gear legs of the main landing gear assembly GAM is arranged. In this way, the motors of the motor assembly E. , also in the spanwise direction of the wing arrangement W. seen, are advantageously installed close to the center of gravity, and because of the smaller moments of inertia, the surfaces of the tail assembly LW can be made smaller to save resistance.

In a further embodiment, at least one motor is the motor assembly E. , in the opposite direction of flight RFD , thus seen approximately in the direction of flow, along the longitudinal axis of the fuselage FA arranged behind at least one bladed rotor R formed rotor system. In this way, the motors of the motor assembly E. in the direction of the longitudinal axis of the fuselage FA advantageously be installed close to the center of gravity, with the surfaces of the tail assembly LW being able to be made smaller to save resistance due to the resulting smaller moments of inertia of the aircraft.

In addition, according to a preferred development of the invention, at least one motor is the motor device E. along the vertical axis VA of the aircraft, in the vertical direction from the ground plane BO seen from below the wing assembly W. arranged. In this way, too, the motors of the motor assembly E. in the direction of the vertical axis VA of the aircraft are advantageously installed close to the center of gravity, with the surfaces of the tail assembly LW being able to be made smaller in a way that saves resistance because of the smaller moments of inertia of the aircraft. A center of gravity that is as low as possible also results in good handling properties of the aircraft in operation flush with the ground and especially when landing.

According to the invention, at least one bladed rotor R is via at least one gear G via at least one motor of the motor arrangement E. driven. It is characteristic of the invention that this driving motor of the motor assembly E. outside the pressurized cabin P. is housed. This is done both for reasons of practicality and for reasons of safety, for example because of the built-in volume, the cooling, the fire hazard and, if necessary, the air supply and exhaust gas discharge. The driving motor of the motor assembly E. can also be wholly or partially in the fuselage arrangement F. , but outside the pressurized cabin P. , be arranged.

In a further preferred development, the gears are also included G for coupling at least one rotor R to at least one motor of the motor arrangement E. also outside the pressurized cabin P. arranged. In a further development of the invention, at least one bearing of the transmission is supported G on the fuselage assembly F. in order to carry away structural bearing forces there. The transmission G can also at least partially within the fuselage arrangement F. run away.

Generally speaking, the gearbox G work mechanically, for example non-positively, positively, etc., also transfer their power hydraulically, electrically or as a suitable mixed form of these conversion types. For example, in a future development, a shaft turbine of the motor arrangement E. drive a suitable generator that provides electrical power for an electric motor that drives at least one bladed rotor R at least in part in terms of power.

According to a further embodiment of the invention, a bladed rotor R can also be preceded or followed by a bladed stator which converts the swirl energy of the air flow at least partially into further propulsion. This could, for example, be arranged in a fixed manner. In a further embodiment, at least individual blades of the stator are designed to be adjustable with regard to the free, undisturbed flow of air, so that they are also advantageous for vectorizing the thrust, for control, for trimming, for torque compensation after an engine failure, to save drag or to increase lift Aircraft can be used with. As with a bladed rotor R, the blades can be adjusted mechanically, electrically, hydraulically or with the aid of air loads via the dynamic pressure.

As a further special aspect of the invention, at least one motor of the motor device E. be designed as a primarily shaft power outputting gas turbine. These have a favorable power-to-weight ratio and operate in a fuel-efficient manner. The shaft power can be transmitted to at least one bladed rotor R, for example, via gears. In addition, they could be kinematically linked to the bladed rotor so advantageously with suitable gears that they work particularly fuel-efficiently for a large part of the flight mission in an advantageous stationary operating mode. This advantage is further enhanced by the invention for the blades B. of the rotor R a blade adjustment is provided for thrust adjustment. In this way, the gas turbine could for the most part work stationary, because the thrust can be adjusted and controlled not only by adjusting the speed of the gas turbine, but also by adjusting the blades.

As a further embodiment, there is the possibility of using at least one motor of the motor arrangement E. designed as a gas turbine, the thermal efficiency of which is further increased in the sense of unconventional measures by recuperative devices for compressor intercooling, fuel mass preheating or exhaust gas heat exchange. As a special consequence of the invention, the space available there, which is at least partially embedded in the aircraft, can be used for devices that are very space-consuming, without the aircraft's resistance increasing significantly as a result. The previous problem with the prior art was that these devices had to be arranged in the immediate vicinity of the engine and thus also in the engine nacelle. However, this would have had to be significantly larger and also disproportionately more resistance-intensive.

According to an advantageous development of the invention, at least one motor is the motor device E. in an area of the aircraft outside the pressurized cabin P. , but at least partially sunk within a space that is exposed to the outside to the fluid through the from the outer skin of the wing assembly W. and that of the outer skin of the fuselage assembly F. formed outer contour of the aircraft is delimited, and is embedded in this space so that a component that acts as an engine mount structural forces from at least one engine of the engine assembly E. in the direction of the structure of the aircraft, is essentially not visible from the outside and that this component in its function of the engine mount outside the, through the from the outer skin of the wing assembly W. and from the skin of the fuselage assembly F. formed outer contour, essentially no separate, resistance-effective surface flushed by the fluid can be assigned.

According to the invention, the engine or engine mounts no longer contribute to the resistance of the aircraft, and the engine nacelle can be dispensed with, apart from air inlet and exhaust gas discharges with their resistance-relevant flushed surface. The efficiency of the aircraft is thus advantageously higher. Due to the embedded position of the engines, engine noise, especially turbine noise, is effectively shielded from the outside.

Thus, in an exemplary embodiment, shaft power turbines or also turboprop engines in the vicinity of the wing / fuselage area, that is to say in the area of the belly fairing, could preferably be partially sunk in an unprinted area U , for example, can be arranged within the landing gear shaft in such a way that the air inlets are suitably far from the unprinted area U can be led out, and also the engine exhaust gases are led out of this. Thus, these motors cause the motor assembly E. , now mostly inside the aircraft, significantly less air resistance, which also reduces the aircraft's fuel consumption again. According to a further aspect of the invention, one or more bladed rotors R can then via long-distance shafts T , which preferably run parallel to the longitudinal axis of the aircraft, or are driven by systems for hydraulic power transmission.

In these exemplary embodiments, the motors of the motor arrangement, which are preferably mounted as shaft drives or turbo propoturbines for driving the bladed rotors E. , sunk and mounted very close to the center of gravity with regard to all coordinate axes. This in turn means for the aircraft, also with the rotors R, which are also attached closer to the center of gravity, that lower moments of inertia result. Above all, however, compared with the previous aircraft configurations with an exclusive rear arrangement of propfan engines, a significantly more flexible loading of the aircraft over a more suitable permissible center of gravity is possible in practice. In addition, the structurally heavier T-tail unit can be dispensed with in favor of a conventional, conventional cross tail unit. At the same time, the external noise of the propeller turbine is greatly reduced by the recessed arrangement in the aircraft. In a further embodiment, the exhaust gas jet could also be led out on the underside of the fuselage rear in such a way that it energizes the boundary layer in the area of the rear fuselage constriction so that aerodynamic detachment is reduced, which could also reduce the resistance of the fuselage. A suitable exhaust system could also shield the jet noise from the engines in the direction of the ground at the same time.

At least one motor of the motor assembly E. was, as shown by way of example, in an unprinted area U of the aircraft housed, for example in a landing gear shaft. However, previously common unprinted areas could also be used U of the aircraft can be expanded for this purpose or new ones are created in addition.

Another possibility and alternative design would be if the motors of the motor assembly E. at least partially in the vicinity of the unprinted rear area of the fuselage arrangement F. are housed. From here they could turn by long-distance waves T or via a system for hydraulic power transmission on the hull assembly F. arranged bladed rotors R drive.

Any desired mixture of the arrangement options described when using several motors and engines would also be conceivable.

With the forms of arrangement shown, the usual, otherwise necessary structural engine connections in the form of pylons, which are known to be quite complex and expensive, are also dispensed with.

An essential inventive aspect is that the bladed rotor R around the fuselage arrangement F. of the aircraft is arranged. This results in the geometric rotor area RAE from the rotation of the bladed rotor around the fuselage assembly F. as a circular ring element with a comparatively large inner diameter ID . This inside diameter ID is also even larger than the outer diameter of previous turbofan engines. In the case of a circular ring element with a comparatively large inside diameter, it is the case that the rotor blades have a certain blade span B. a much larger geometrical rotor area RAE can be achieved than with a circular element with the same blade span. The largest possible rotor area, however, is the prerequisite for a particularly fuel-efficient propulsion of the aircraft with a low FPR and high mass flow. According to the invention, in a further advantageous embodiment, the smallest inner diameter is ID at least one bladed rotor R selected from the dimensions at least larger than the vertical height V the pressurized cabin P. and at the same time the vertical height V this pressurized cabin P. within the fuselage assembly F. designed in such a way that an upright gait is possible there for the average passenger, at least in sections. Thus, on the one hand, there is the advantage that the passengers are in the pressurized cabin P. can stand and walk comfortably, on the other hand, according to the invention, the geometry for the bladed rotor R is given at the same time in such a way that a large geometric rotor area is present RAE with a small required span of the rotor blades B. sets and thus sets a particularly efficient generation of propulsion and fuel efficiency on the aircraft.

In addition, the smaller the radial span of the blades required B. Geometric framework conditions that primarily concern the maximum radial blade extension on the aircraft, for example with regard to the ground clearance of the aircraft or the available height between the wing and the floor, are no longer so important according to the invention.

By inventive achievement of a large geometrical rotor area RAE of the bladed rotor R with a comparatively small radial blade extension of the blades B. According to the invention, a circular ring element with a large inner diameter results in a blade under consideration B. of the rotor R a reduced difference in the bandwidth of the circumferential speeds occurring in the radial direction depending on the span. This favors the aerodynamic design of the bladed rotor to achieve a high degree of component efficiency and therefore contributes significantly to the overall efficiency of the aircraft. As a result, less twisting could be necessary over the blade span to generate a certain thrust, which increases the aerodynamic blade quality and helps to reduce the blade weight and its complexity. Likewise, a blade adjustment can be realized in a technically simpler manner through this and, in particular, through the lower extension on the side of the span, which also reduces the blade moment at the point of mounting.

With the small radial speed difference according to the invention in the span-dependent flow onto the rotor blades B. Overall, according to the invention, it is easier to choose an adapted, noise-optimized angular speed of rotation for the bladed rotor R. In particular, it would then be possible to be able to select this rotation speed so that it is less likely that the outer blades will get into such an inflow state, for example a critical or supercritical one that generates disproportionately strong noise from the outside perception. A high aerodynamic efficiency of the bladed rotor R according to the jet theory with as uniform and homogeneous a speed distribution as possible over the entire surface of rotation of the bladed rotor R can thus be achieved more easily according to the invention.

In addition, according to the invention, the rotational speed of the rotor can preferably be optimally matched to it. The transmission G in their transmission ratio in the drive train can be selected so that not only for the bladed rotor R, but also for at least one motor of the motor arrangement E. optimum operating conditions in terms of rotation speed and moment exist, which increases the overall efficiency of the drive system. By tuning to the bladed rotor R, the speed of rotation could be with the enlarged thrust cross-sectional area or rotor area RAE for example, chosen to be relatively low. The thrust is then mainly achieved via a large mass flow, the blades B. of the rotor R run relatively slowly at their hub as well as at their tip, which keeps the noise low.

At the same time, the thrust generation of the bladed rotor R is concentrated according to the invention by the circular ring element on the radial section around 70-75% of the rotor diameter, which can generally generate thrust most effectively and efficiently in rotors.

This is not the case with previous prior art engines. If, for example, in a conventional turbofan or propeller engine, the outer blade boundary rotates very close to the critical state at a certain angular speed or is already locally supersonic, the opposite end of this blade near the hub moves at the same angular speed due to the large necessary extension of the blade to achieve a certain thrust area, but with a very much reduced circumferential speed, which even reaches zero directly at the hub. The difference in the circumferential and thus also in the flow velocity of the blades is therefore very large in the previous bladed rotors of conventional engines, disadvantageously reducing the efficiency.

In addition, if desired, the thrust cross-sectional area of the bladed rotor R, which can be increased according to the invention, can increase the bypass flow ratio, which increases the efficiency of the drive system and further lowers the noise and increases the propulsion efficiency of the aircraft and thus its efficiency.

In a further preferred embodiment of the invention, according to the invention, that of the rotor blades, which is the measure for the generation of propulsion, can be used B. of a bladed rotor R in rotation swept over the geometric rotor surface RAE be designed to be the same size or larger than the cross-sectional area of the fuselage arrangement that is considered a measure of the fuselage resistance of the flight and is enclosed by the bladed rotor R F. , which is in one of the rotor levels RP this bladed rotor R perpendicular to its geometric axis of rotation RA results.

According to the invention, it is thus possible to have a rotor centrally on the fuselage arrangement F. to be arranged in its rotor surface RAE , even exceeds an area which in the cross-section of the fuselage assembly F. just with the diameter of this fuselage arrangement F. can be formed. According to the prior art, it would have hitherto only been possible to achieve the cross-sectional area of the fuselage with a bladed rotor R by conventionally selecting the rotor diameter equal to or greater than the fuselage diameter via a circular element and this rotor next to and outside the fuselage arrangement F. is arranged. However, this rotor must be in a plane perpendicular to the longitudinal axis of the fuselage FA at least half a fuselage diameter distance from the outer boundary of the fuselage arrangement F. of the aircraft so that it can rotate freely. Commercial aircraft require redundancy with at least one second rotor for safety reasons in the event of a rotor failure. If, however, one of these two rotors, which are adjacent to the fuselage arrangement, fails, a significant disadvantageous moment results from the lever arm the size of at least one fuselage diameter. Due to the now central arrangement according to the invention on the fuselage arrangement F. this moment can now be avoided and a high efficiency of the drive can still be achieved over a large rotor area.

In addition, in a further advantageous embodiment, that of the rotor blades, which is the measure for the generation of propulsion, can be used B. of a bladed rotor R in rotation swept over the geometric rotor surface RAE be made the same size or larger than fifteen times the sum of the air inlet cross-sectional areas of all air-breathing motors of the motor assembly E. that drive this bladed rotor R, the rotor blades B. of this bladed rotor R are designed aerodynamically so that the aircraft can reach flight speeds of at least Ma 0.78 at altitudes greater than 10000 m. In this way, it is possible with the drive arrangement to achieve higher bypass flow ratios in practice than is possible today with geared turbofan drive units that achieve bypass flow ratios of a maximum of 15: 1 or less. At the same time, the cruising speeds and altitudes that are higher for turbofan engines can also be achieved, which are overall greater than for turboprop-powered aircraft.

The advantages of the arrangement described also enable the use of open rotors and propfan rotors and drive systems on the aircraft in a safe manner. This means that higher bypass flow ratios can also be achieved at the comparatively high cruising speeds of today's turbo engines. This significantly reduces fuel consumption compared to aircraft in operation today.

In general, the inventive arrangement of the rotor, ring-shaped surrounding the fuselage, results in the possibility of using more freedom in the aerodynamic design of the bladed drive rotor and, retrospectively, also of the drive system, which overall favor a more fuel-efficient and quieter design of the drive system as well as a higher propulsion efficiency enable the aircraft.

Thus, in an exemplary embodiment, the effective circular thrust area could be selected to be larger overall in terms of area. For a certain airspeed, in order to achieve a predetermined thrust, there is then the possibility of being able to reduce the speed distribution given to the fluid by the bladed rotor R and thus to be able to lower the pressure ratio of the bladed rotor R. In this way, the load on the individual blades of the rotor can also be reduced, for example by reducing the angle of attack or setting angle of the blades, or by changing the profile to produce less aerodynamic lift per length. As a consequence, there is a lower thrust circular area loading, or in this case a lower thrust circular ring area loading, less rotor noise and increased fuel efficiency.

In a further embodiment, however, if the remaining parameters such as the pressure ratio of the bladed rotor R, the angular speed of the rotation and the pitch of the blading, based on a reference engine, are left approximately the same with a larger thrust cross-sectional area, the result is a higher thrust for the aircraft.

With a constant thrust cross-sectional area compared to a current reference engine, in a further possible embodiment with the same pressure ratio of the bladed rotor R there is the possibility of more rotor blades B. to be arranged on the rotor surface in order to achieve a lower necessary angular speed in the rotation, with the consequence of a significantly smaller difference in the circumferential speed between the root and tip of a rotor blades B. which primarily reduces noise and can also increase efficiency.

In a further exemplary embodiment, this is done with the same thrust cross-sectional area and the same circumferential speed at the outer tips of the rotor blades B. by regulating the angular speed of rotation set to approximately the same circumferential speed as with a current engine, this results in a higher output thrust with otherwise unchanged aerodynamic design parameters and approximately the same expected noise effect.

With a variable setting angle of the blades B. of the rotor R, the blades of the rotor R can, if necessary, even be brought into the sail position, for example in the zero lift direction, whereby the additional drag can be significantly reduced by this feathering. The adjustment can also be organized automatically in the event of an engine failure. It is another important performance advantage compared to conventional engines. In a further embodiment, the blades can also be folded in at the rear, for example for a fuselage arrangement F. be brought into the sail position, similar to a folding propeller. It is precisely this that would be of significant performance advantage in multi-engine aircraft after an engine failure.

Another possibility could be the adjustment of the rotor blades B. , for example in their setting angle, also be designed so that a thrust reversal effect of the bladed rotor R for braking on the ground and possibly also in the air is possible. In the event of a malfunction, there is advantageously no moment which the aircraft would like to take away from the runway.

The bladed rotor R, also often referred to as a “propulsor” in its mode of operation, is designed in such a way that, as explained above, it is supported by a suitable bearing arrangement S. can rotate on the aircraft in relation to this and through a suitable design of the blades B. is technically suitable for converting mechanical drive power into propulsion power of the aircraft. The bladed rotor thus generates thrust with the effect of a propulsive force on the aircraft. This propulsive force driving the aircraft is oriented in such a way that its force component is in the direction of the longitudinal axis of the fuselage FA compared to the force component perpendicular to the longitudinal axis of the fuselage FA clearly predominates. The bladed rotor R takes over the generation of propulsion in an essentially horizontal direction of flight and thus essentially does not contribute, according to its function, to generating the decisive lift for the aircraft, as in the case of a rotary wing aircraft.

With regard to its blading, the bladed rotor can, in one possible embodiment, be designed in such a way that the blades B. of the rotor R with regard to characteristic, aerodynamic parameters, for example the thrust cross-sectional area, circular (ring) area load, rotational angular speed, blade span, blade depth, relative thickness of the blade, twisting, tapering, blade sweep, profile, the number of blades and the overall appearance are designed similar to the blades of a fan .

In a further embodiment, the bladed rotor R can be designed in such a way that the blade blades B. of the rotor with regard to characteristic, aerodynamic parameters, for example the thrust cross-sectional area, circular (ring) area loading, rotational angular speed, blade span, blade depth, relative thickness of the blade, torsion, tapering, Blade arrowhead, profiling, the number of blades and the overall appearance are designed similar to the blades of a prop fan.

In addition, in a further embodiment, the bladed rotor R can be designed in such a way that the blades ( B. ) of the rotor with regard to characteristic, aerodynamic parameters, for example the thrust cross-sectional area, circular (ring) area load, rotational angular speed, blade span, blade depth, relative thickness of the blade, twisting, tapering, blade sweep, profile, the number of blades and the overall appearance are designed similar to the blades of a propeller .

In addition, any advantageous mixtures of the aforementioned embodiments are possible with regard to the characteristic design parameters.

In an advantageous further embodiment, the blades are B. in an outward radial extension, based on suitable characteristic, aerodynamic design parameters, so that they can have less of a propulsion-generating effect at a span position near their hub than at a span position that is about 2/3 of the respective maximum blade span. This can be done, for example, via the span-dependent profiling, also via a twisting of the rotor blades B. happen. It has the advantage that the propulsion generation takes place outside the hull boundary layer and thus possibly more effectively and with less noise. This does not rule out that the hull boundary layer is also moved or accelerated with the aid of a suitable span-dependent design of the blades. The blading in this span section could advantageously be designed according to a type of desired speed profile of the boundary layer.

According to a further embodiment, rotor blades could also be used B. be designed differently from one another, for example have a different span in order to influence the noise favorably. In rotary operation, this nevertheless results in an annular thrust cross-sectional area of the bladed rotor.

For the execution of the bladed rotor R it could also be useful in a further embodiment to select a very specific even or odd number of blades in order to keep the noise even lower, for example by limiting certain natural oscillation frequencies.

According to the invention, the bladed rotor R is also suitable for being able to be kinematically connected to a motor arrangement of the aircraft by suitable features as part of a transmission so that the rotor R can generate propulsion driving the aircraft through its rotation.

In a preferred embodiment, it can be coupled to the engine arrangement of the aircraft as part of a mechanical transmission. For this purpose, it has, for example, a suitable toothing along one of its sides or along its circumference as a feature. This toothing can preferably be arranged on the inside or on the outside, the sides or also on any inclined plane and also use a toothing shape that ensures a reliably high power transmission, for example a helical toothing or a toothing on which bevel gears can engage. The rotor is then for example for driving at least one of at least one motor of the motor arrangement E. The bladed rotor could, in a further embodiment, also be designed as part of a hydraulic transmission. This transmission should have the highest possible efficiency.

If it coincides at least partially with an annular electric motor, it thereby forms part of an electrical, sometimes also an electronic, transmission.

According to a particular aspect of the invention, bladed rotors R and drive systems rotating in opposite directions can also be used by different power control to control the aircraft around the longitudinal axis while rolling. The at least partial roll control or trimming could take place via different speeds, torque specifications or by different positions of the blades of the rotors and possibly the stators.

According to a particular development of the invention, the bladed rotors R can be driven by one or more motors of the motor arrangement E. over long-distance waves T or systems for hydraulic power transmission are driven in the unprinted areas U are at least partially housed within the aircraft.

The bearing arrangement S. can preferably be done by a form of arrangement of roller bearings, plain bearings or a suitable combination of both types of bearings. In a further embodiment, the bearing arrangement could S. of the rotors can also be formed by hydraulic bearings, for example hydrostatic bearings. Furthermore, magnetic bearings could also be used to support the rotor. Both types of bearings mentioned above offer the possibility, also through active control, the noise and to be able to additionally dampen the vibrations towards the aircraft. In a further possible arrangement, piezoelectrically variable and electrically controllable dampers could be used within the bearing from the bladed rotor to the fuselage, which allow active oscillation and vibration damping.

Magnetic bearings also ensure extremely low bearing friction, which has a beneficial effect on the efficiency of the drive system. According to the current state of the art, magnetic bearings, e.g. superconducting magnetic bearings or electrodynamic magnetic bearings, require a certain amount of installation space, which can easily be set up by the inventive arrangement of the bladed rotor on the outside of the fuselage in combination with the large support width in the bearing.

The drive unit, formed from the bladed rotor R and the associated bearing and guide unit, can also have panels O , e.g. also for the fuselage arrangement F. of the aircraft in order to additionally influence the flow aerodynamically or to increase the aerodynamic quality of the drive system.

In a further embodiment, the mounting can optionally also take place together with an electric motor, for example in such a way that the bladed rotor R advantageously also forms the rotor of the electric motor at the same time. The mounting of the common rotor can also be designed in such a way that the mounting and electric motor share other component components, for example magnets, coils or a cooling system.

In a further embodiment, a rotor R could also be part of a “stator” of at least one electric motor and another rotor R could form part of a rotor of an electric motor so that these rotors move in opposite directions to one another and do not apply a moment to the fuselage arrangement F. is transmitted.

Since engine mounts and engine nacelles no longer have to be accommodated on the wing, the aerodynamically clean wing of the wing arrangement can now be used W. preferably be equipped with additional concepts to save aerodynamic drag, for example with laminar maintenance technologies, such as the NLF-Wing (Natural Laminar Flow Wing). Furthermore, the integration of a high lift system on the wing without engines is now generally easier and technically more effective.

An advantage according to the invention is that the bladed rotor R is rotatably mounted with little additional weight, preferably structurally directly on the fuselage arrangement F. can be connected.

In this sense, no separate engine mounting concept is necessary, but a reinforcement of the existing fuselage structure could be sufficient here, whereby in the sense of lightweight construction the high polar moments of inertia and the high torsional moment of resistance of the fuselage arrangement, due to the opposite fuselage side walls, are particularly advantageous can be used.

Due to the large support width in the bearing of the bladed rotor R, due to the opposite side walls of the fuselage, approximately in the order of magnitude of the outer diameter of the fuselage, a stable guidance of the bladed rotor R results in rotary operation. During operation, this leads to a comparatively insensitivity to a rotor R which is not completely correctly mechanically or aerodynamically balanced and is therefore out of true running bladed rotor R, for example after damage to individual blades B. , for example due to foreign bodies in operation, would be conceivable. In terms of design, operation can thus also be possible after damage, ie a fault-tolerant design of at least one bladed rotor R appears to be possible in a further embodiment.

Both due to the greater support width in the bearing and due to the smaller blade extension of the blades according to the invention B. This also results in a significantly reduced risk of vibrations, both during normal operation and after damage to the rotor R. This eliminates another problem with previous propan and open rotor drive arrangements, namely the possible occurrence of increased vibrations coupled with noise, improved.

According to a further aspect of the invention, multiple motors of the motor arrangement E. drive a bladed rotor R in parallel. With regard to the possibility of damage-tolerant design of the bladed rotor R, it would also be conceivable to keep the rotor in fail-safe design on the aircraft and thus to meet the redundancy requirements of the drive system by providing several motors that could only drive one rotor in parallel R. The architecture can also be selected so that in the event that one of the motors of the motor arrangement E. fails, the remaining motor or motors continue to provide sufficient drive power to drive the bladed rotor R so even in critical flight phases of the aircraft that its effect does justice to the performance requirements of the aircraft.

According to a further advantageous aspect of the invention, at least one of the motors of the motor arrangement could E. can be optionally coupled in or out by a coupling for driving the rotor R. In the event of a malfunction of a motor, for example, it could be decoupled from the drive system. If the power requirement increases, a further motor could preferably be coupled in. Likewise, the rotors on the floor can also be decoupled if necessary, sometimes also for rolling.

According to a further advantageous aspect of the invention, at least one of the motors of the motor arrangement E. not only be designed as an internal combustion engine, for example as a gas turbine or piston engine, but also as an electric motor.

This results in the basic advantages of an electric drive in an order of magnitude, corresponding to the proportion of the electrical power used in relation to the total power. In the case of a partially electric drive in the sense of a hybrid concept or fully electric drive, the advantages of the electric drive come into play, such as minimal maintenance, whereby many components are even wear-free, better adaptability of torque and speed to the bladed rotor, advantageous efficiency, a high efficiency of the order of 96% for electric motors, compared to 36% for gas turbines, a high ratio of the maximum to continuous output of the electric motor of 2 - 2.5, easy replacement of components, low complexity, lower acquisition costs for the electric motor with the possibility of lower overall operating costs.

In addition, there is the further advantage that motors of the motor arrangement E. , which are accommodated, for example, in the area of the wing-fuselage arrangement or in the landing gear shaft, are very easily accessible from the outside for maintenance, repair or replacement purposes. This applies regardless of whether the drive system is conventional or hybrid-electric, and also for the other components of the drive system, such as for the bladed rotors R and possibly for the long-distance shafts T which can be attached, for example, underneath an unprinted aerodynamic cladding on the outside of the fuselage, lying on this.

However, there is also the special advantage that aircraft with this drive system can initially be designed conventionally as internal combustion engines. With increasing success in electromobility research, an engine could only be replaced by an electric motor, resulting in a hybrid-electric powered aircraft. The other components of the drive system, in particular the bladed rotor R and its kinematic coupling to the motors, could probably remain unchanged. With further advances in research, additional engines could possibly be replaced by electric motors, so that in the end a fully electrically powered aircraft would be available with a manageable technical and low investment risk. At the same time, attention could be paid to a universal engine mount right from the initial design on the aircraft, which would enable the later replacement of internal combustion engines with electric motors.

In addition, in a further embodiment, at least one electric motor of the motor arrangement can be used as part of a hybrid electrical system E. also temporarily at least temporarily provide additional propulsion power, preferably in flight phases that require at least a short time an increased thrust requirement of the aircraft, for example during take-off, climbing or in the event of an engine failure.

According to a further advantageous aspect of the invention, an annular electric motor could also be part of the motor arrangement E. , essentially concentric to the bladed rotor R, the fuselage F. surrounding, be arranged to drive the bladed rotor R. This electric motor could also be partially embedded in the fuselage arrangement or also a component of the fuselage arrangement F. form. Powerful ring-shaped electric motors are known from the generation of wind energy, which are used there as generators and which in principle could also be operated as motors. Current research goes assume that the mass-specific power-to-weight ratio of these components can be improved by a significant factor in the future.

This type of arrangement also offers the advantage that an electric motor, which is in a kinematic alternating connection with the bladed rotor, is effectively cooled during operation by the outside environment, which is about -60 ° C cold when cruising.

In general, it would also be possible to use electric motor types suitable in the future, such as superconducting motors, for example HTS motors or also transverse flux machines within the motor arrangement E. conceivable.

According to a further advantageous embodiment of the invention, at least one air-breathing motor of the motor arrangement that drives at least one bladed rotor R generates, in addition to the propulsive force of at least one rotor R, further propulsive force in the form of thrust through the ejection of exhaust gases, the force component in Longitudinal direction of the axis of rotation RA compared to the force component perpendicular to the axis of rotation RA clearly predominates. In this way, according to the invention, the bladed rotor R is relieved of its area-specific thrust load, so that it can work more effectively in certain operating states of the aircraft, for example when taking off. Similar to a turbofan engine, this results in an engine arrangement at aircraft level, the bypass flow enveloping the fuselage arrangement and the shaft turbine with the core flow also generating thrust, but at a larger dimensional level and also at the level of the possible bypass flow ratio.

It would also be possible to use this residual thrust of at least one air-absorbing motor to roll the aircraft on the ground. For this purpose, the rotor R could also be temporarily decoupled from this motor.

In a further embodiment, the largest outside diameter is OD at least one bladed rotor is designed to be equal to or smaller than twice the maximum longitudinal extension LGAM a landing gear leg with wheels WE the main landing gear arrangement GAM . This is an arrangement of the main landing gear GAM possible in such a way that the two landing gear legs attached to the wings neither interfere with the inflow nor the outflow of a rotor R. Likewise, as indicated here, the chassis can be pivoted inward and retracted. Due to this geometric arrangement, the available space is also sufficient in length for the retraction of both landing gear legs. At the same time, the rotor cannot geometrically touch the ground plane.

According to an advantageous development of the invention, at least one bladed rotor R becomes the fuselage arrangement F. At least partially surrounding diffuser is arranged, which advantageously reduces the speed of the fluid additionally induced by the bladed rotors and at least partially converts it into additional pressure, which, due to the curvature of the surface, has a thrust-generating effect on the aircraft due to the relative overpressure in the flow (similar to the principle of thrust recovery). This diffuser can also be changed by changing the fuselage diameter of the fuselage assembly F. or given by their disguises.

If an electric motor is used in the drive system which is in kinematic alternating connection with the bladed rotor R, this could preferably also be operated as a generator in a further embodiment, e.g. in suitable operating phases of the aircraft.

With these aspects summarized in their effect, the invention thus lays the foundation for realistic architectures of future hybrid-electric propulsion systems in aircraft.

On the ground, it could also be advantageous not to use one of the drive systems according to the invention for taxiing on the apron, but to use an electric wheel drive system for this, in particular also at the parking position. In that the thrusters of the engine assembly E. In possible embodiments of the invention can be separated from the rotor R by clutches, if internal combustion engines are also used, starting and warming up during operation, for example the gas turbines, is also possible without rotating the rotor R. For this purpose, centrifugal clutches could also be used. This means that ground handling employees cannot be endangered by the operation of the rotors.

In a further embodiment, the bladed rotor R can also be used to accelerate the fuselage boundary layer, ie it supplies energy, which can thus additionally reduce the resistance of the fuselage and the aircraft. In a further embodiment, the fuselage boundary layer could at least partially be suctioned off at least partially upstream of at least one rotor R in order to avoid an unfavorable interaction with the rotor R.

According to an advantageous development of the invention, the nose landing gear arrangement ATM regardless of the main landing gear arrangement GAM can be retracted and extended again. Thus, according to the invention, bladed rotors R can forward over the fuselage arrangement F. can be withdrawn for replacement, maintenance or servicing without having to be dismantled in their scope.

It should be noted that within this document, the thrust cross-sectional area or the cross-sectional area of the bladed rotor in the sense of a rotor circular area or a rotor circular ring area is understood to mean that area of the bladed rotor which, during rotary operation, is caused by sweeping over the blades B. arises in a plane normal to the axis of rotation RA of the rotor R is stopped.

It should also be noted that the term turbo-propeller in this document refers to the function of the gas turbine unit generating the shaft power without a propeller. In one possible embodiment, this can already have a reduction gear internally.

The exemplary applications on aircraft configurations shown in the document are to be understood as exemplary and include the application of the invention also on other aircraft configurations with differing features. e.g. with regard to the tail unit or engine arrangement.

RFD denotes a direction that is opposite to the flight direction of the aircraft, i.e. corresponds approximately to the direction of flow of the fluid.

The patent application seeks protection for a fixed-wing aircraft that does not take off and land vertically, preferably of the type of a commercial aircraft.

In addition, it should be noted that “comprehensive” does not exclude any other elements or steps and “one” or “one” does not exclude a plurality. It should also be pointed out that features or steps that have been described with reference to one of the above exemplary embodiments can also be used in combination. Reference signs in the claims are not to be regarded as a restriction.

Abbreviations BPR By pass ratio CRTF Counter rotating turbo fan FPR Fan Pressure Ratio HTS High Temperature Superconductor (high temperature superconductor) NLF Natural laminar flow UDF Unducted fan UHB Ultra high bybass

List of reference symbols

B.

Rotor blade

BO

Ground level

E.

Motor arrangement (formed by at least one motor)

F.

at least partially cylindrical fuselage arrangement

FR

Flight direction

FA

Longitudinal axis of the trunk

G

transmission

GA

Landing gear arrangement

ATM

Nose gear assembly

GAM

Main landing gear arrangement

ID

smallest inner diameter of a bladed rotor

LGAM

Largest length of a main landing gear leg with wheels WE of the main landing gear arrangement GAM

MV

Minimum vertical level of the upper side of the wing arrangement W in the direction of the vertical axis of the aircraft VA as seen from the ground plane BO

O

Disguise

OD

largest outside diameter of a bladed rotor

P.

in the function of a pressurized cabin, printable area within the fuselage arrangement (pressurized cabin), in another A.-Form: pressurized cabin with overpressure

RA

geometric axis of rotation of a bladed rotor

RAE

Geometric rotor surface swept over by the revolution of the rotor blades B.

RFD

Direction opposite to the direction of flight

RP

Rotor plane, which is within the longitudinal extent of a rotor R, along its axis of rotation RA normal to this axis of rotation RA, as an auxiliary plane, with an extent greater than the spatial extent of the flight, for determining the area of impact of the bladed rotor R.

S.

Bearing arrangement

T

Long-distance wave

U

Unprinted area of the aircraft

V

Vertical height of the cabin within the fuselage arrangement

VA

Vertical axis of the aircraft

W.

Wing arrangement

WE

wheel

Figure list

• 1 eine stark vereinfachte beispielhafte Darstellung der Seitenansicht eines zukünftigen Verkehrsflugzeuges mit zwei drehbar gelagerten beschaufelten Rotoren R, wobei jeder Rotor R durch jeweils eine zugeordnete Wellenturbine der Motorenanordnung E über eine zugehörige Fernwelle T rotatorisch angetrieben werden kann und, wobei sich die Wellenturbinen innerhalb eines unbedruckten Bereiches U in der Nähe der Flügel-Rumpf-Anordnung, dort teilweise im Flugzeug innen liegend versenkt, angeordnet befinden, und wobei in dieser Darstellung die zweite Wellenturbine mit ihrer Fernwelle nicht direkt ersichtlich ist, da sie, bedingt durch die Darstellung, hinter dem ersten Antriebssystem in Spannweitenrichtung verdeckt angeordnet ist;
• 2 eine stark vereinfachte beispielhafte Darstellung der Vorderansicht eines zukünftigen erfindungsgemäßen Verkehrsflugzeuges mit zwei drehbar gelagerten beschaufelten Rotoren R, wobei jeder Rotor R durch jeweils eine zugeordnete Wellenturbine der Motorenanordnung E über eine zugehörige Fernwelle T rotatorisch über Getriebe angetrieben werden kann und, wobei sich die Wellenturbinen innerhalb eines unbedruckten Bereiches U in der Nähe der Flügel-Rumpf-Anordnung, dort teilweise im Flugzeug innen liegend versenkt, angeordnet befinden;
• 3 eine Übersicht bisher üblicher bedruckter Bereiche P und unbedruckter Bereiche U innerhalb der Rumpfanordnung F eines Transportflugzeuges;
• 4 die beispielhafte sichere Anordnung von zwei beschaufelten Rotoren R entlang der Rumpf-anordnung F am Flugzeug entsprechend einer alternativen Ermittlungsmethode so, dass sich innerhalb der, mittels virtueller Torus-Körper symbolisierten Ausflugs-Kegeln mit gewünschten geeigneten Öffnungswinkeln, die von den beschaufelten Rotoren ausgehen, keine für den sicheren Flug notwendigen weiteren Komponenten des Flugzeuges befinden;
• 5 eine stark vereinfachte beispielhafte drehbare Anordnung eines erfindungsgemäßen Rotors R mit einer Lageranordnung S, den Profilschnitt der Rumpfanordnung F ringförmig umgebend, wobei innerhalb des Querschnittes der Rumpfanordnung F ein gemäß einer Druckkabine P bedruckbarer Abschnitt vorgesehen ist, der von seiner vertikalen Bauhöhe V so hoch ausgeführt ist, dass ein durchschnittlicher Fluggast in ihr aufrecht stehen kann;
• 6 Profilschnitt einer beispielhaften Ausführung einer Lager- und Führungseinheit mit einer Lageranordnung S zur rotatorischen Lagerung des Antriebsrotors R bei gleichzeitiger antriebswirksamer Abführung der Axialkräfte und kinematischer Ankuppelung des beschaufelten Rotors R über Getriebe an eine Fernwelle T,
• 7 eine stark vereinfachte beispielhafte kinematische Ankoppelung eines erfindungsgemäßen beschaufelten Rotors R über Getriebe an mehrere Motoren der Motorenanordnung E, wobei der beschaufelte Rotor R eine Innenverzahnung aufweist, auf die ein weiteres Zahnrad, das mit einem Motor der Motorenanordnung E verbunden ist, kinematisch einwirkt, hier beispielhaft als hybrid-elektrische Ausführungsform des Antriebssystems, wobei ein Motor der Motorenanordnung E über eine Kupplung ein- und auskuppelbar ausgeführt ist; wobei ein Motor als Gasturbine und der weitere Motor als Elektromotor ausgeführt ist;
• 8 zeigt eine stark vereinfachte beispielhaft skizzierte Darstellung der möglichen Anordnung eines beschaufelten Rotors, gezeigt in einer Ebene an einer Rumpflängenposition in Höhe der Tragflächenanordnung W stromabwärts der Hauptfahrwerksanordnung GAM senkrecht zur Rotationsachse eines beschaufelten Rotors, wobei der größte Außendurchmesser OD wenigstens eines beschaufelten Rotors gleich oder kleiner ausgeführt ist als die doppelte maximale Längserstreckung eines Fahrwerksbeines mitsamt Rädern WE der Hauptfahrwerksanordnung GAM;
• 9 zeigt eine stark vereinfachte beispielhaft skizzierte Darstellung eines mit Rotorblättern B beschaufelten Rotors R, links in der Seitenansicht und rechts in der Vorderansicht, und verdeutlicht die Definition von Rotorebenen RP im Bereich der Längserstreckung des beschaufelten Rotors R entlang seiner Rotationsachse RA, wobei diese Rotorebenen RP zu der Rotationsachse RA normal stehen;

Further details and features of the invention emerge from the following description in conjunction with the drawing. It shows:

• 1 a greatly simplified exemplary illustration of the side view of a future commercial aircraft with two rotatably mounted bladed rotors R, each rotor R being driven by an associated shaft turbine of the motor arrangement E. via an associated long-distance wave T Can be driven in rotation and, with the shaft turbines within an unprinted area U are located in the vicinity of the wing-fuselage arrangement, there partially sunk inside the aircraft, and the second shaft turbine with its long-range shaft is not directly visible in this illustration, because it is behind the first drive system in Span direction is concealed;
• 2 a greatly simplified exemplary representation of the front view of a future commercial aircraft according to the invention with two rotatably mounted bladed rotors R, each rotor R being driven by an associated shaft turbine of the motor arrangement E. via an associated long-distance wave T Can be rotationally driven via gearbox and, with the shaft turbines within an unprinted area U are located in the vicinity of the wing-fuselage arrangement, partially sunk there in the aircraft lying on the inside;
• 3 an overview of the previously common printed areas P. and unprinted areas U within the fuselage assembly F. a transport aircraft;
• 4th the exemplary safe arrangement of two bladed rotors R along the fuselage arrangement F. on the aircraft according to an alternative determination method in such a way that within the Excursion cones symbolized by means of virtual torus bodies with the desired suitable opening angles, which start from the bladed rotors, no further components of the aircraft necessary for safe flight are located;
• 5 a greatly simplified exemplary rotatable arrangement of a rotor R according to the invention with a bearing arrangement S. , the profile section of the fuselage arrangement F. surrounding ring-shaped, being within the cross-section of the fuselage assembly F. one according to a pressurized cabin P. printable section is provided which depends on its vertical height V is designed so high that an average passenger can stand upright in it;
• 6th Profile section of an exemplary embodiment of a bearing and guide unit with a bearing arrangement S. for the rotary mounting of the drive rotor R with simultaneous drive-effective dissipation of the axial forces and kinematic coupling of the bladed rotor R via a gearbox to a remote shaft T ,
• 7th a greatly simplified exemplary kinematic coupling of a bladed rotor R according to the invention via gears to several motors of the motor arrangement E. , wherein the bladed rotor R has an internal toothing on which a further gear, which is connected to a motor of the motor assembly E. is connected, acts kinematically, here by way of example as a hybrid-electric embodiment of the drive system, with a motor of the motor arrangement E. is designed to be engaged and disengaged via a clutch; one motor being designed as a gas turbine and the further motor being designed as an electric motor;
• 8th shows a greatly simplified exemplary sketched representation of the possible arrangement of a bladed rotor, shown in a plane at a fuselage length position at the level of the wing arrangement W. downstream of the main landing gear assembly GAM perpendicular to the axis of rotation of a bladed rotor, with the largest outside diameter OD at least one bladed rotor is designed to be equal to or smaller than twice the maximum longitudinal extension of a chassis leg including the wheels WE the main landing gear arrangement GAM ;
• 9 shows a greatly simplified exemplary sketched representation of a with rotor blades B. bladed rotor R, on the left in the side view and on the right in the front view, and clarifies the definition of rotor planes RP in the area of the longitudinal extent of the bladed rotor R along its axis of rotation RA , these rotor planes RP to the axis of rotation RA stand normally;

Detailed description of the figures

1 The figure shows, in a greatly simplified form, an exemplary embodiment of a future commercial aircraft according to the invention in a side view. According to an important aspect of the invention, two bladed rotors R are a fuselage arrangement that can be printed at least in sections F. , this ring-shaped surrounding rotatably mounted, arranged that each rotor R in one of its rotor planes RP , which in the area of its length along its rotor axis RA are arranged perpendicular to this, at least part of a pressure cabin P. that is inside the fuselage arrangement F. lies, surrounds. According to the invention, this has the consequence that in the event of a rotor breakage during operation, fragments of the rotor R move away from the fuselage arrangement F. , but especially the pressurized cabin P. , remove by the effective centrifugal forces. According to the invention, each rotor surrounds R 1 in one of its rotor planes RP seen, here that cross section of the fuselage assembly F. , which in this plane at least in sections according to a pressure cabin P. is printable. The bladed rotors R are rotatable in relation to the fuselage arrangement F. by at least one bearing arrangement S. (as shown in ) stored and approximately in the direction of the longitudinal axis of the fuselage FA by at least one bearing arrangement S. (as shown in ) fixed arranged. In this exemplary embodiment, the axes of rotation are RA , which coincidentally coincide here with these two rotors R, in the sense of a usual drive camber at a small angle in relation to the longitudinal axis of the fuselage FA run inclined. The two rotors R are inventive along the longitudinal axis of the fuselage FA on the fuselage assembly F. arranged that in all of the rotor planes RP of each of the two rotors that are perpendicular to the axis of rotation RA of the bladed rotor R and are arranged within its longitudinal extent, no further important safety-critical components necessary for the immediately safe flight of the aircraft are arranged. Thus, in the event of a bladed rotor R breaking, these cannot be damaged in a safety-critical manner. Also here, according to a special characteristic of the invention, both bladed rotors R are designed in such a way that they rotate freely in the fluid without a jacket. As a result, they achieve a very high level of efficiency in the propulsion conversion. In this application example, the two bladed rotors R are advantageously installed in such a way that they rotate in opposite directions to one another. Together they form a rotor system. In this way the rotor efficiency of the rotor system is improved. The bladed rotors R are on the fuselage assembly F. these are not only arranged in a surrounding manner, but in a certain way compared to a 3-point chassis arrangement GA , overall with several rotatably mounted wheels WE , and with a nose landing gear ATM , in which GA with the statically determined support of the aircraft against the ground level BO serves in operation flush with the ground, positioned so that both bladed rotors R on the fuselage assembly along the longitudinal axis of the fuselage FA between a nose landing gear structurally attached to the fuselage assembly ATM and a main landing gear assembly GAM are positioned, the latter being attached to the wing assembly W. is structurally connected and includes two main landing gear legs. Since the rotation during take-off and the derotation during landing are essentially centered around the wheels of the main landing gear assembly GAM takes place, there is therefore, according to the invention, no risk of a bladed rotor R on the ground plane even during these critical phases BO touches down or its ground clearance is impermissibly reduced during the rotation or derotation of the aircraft. Corresponding to this design, the dimension for the generation of propulsion is that of the rotor blades B. of a bladed rotor R in rotation swept over the geometric rotor surface RAE designed to be the same size or larger than the cross-sectional area of the fuselage arrangement that is considered a measure of the fuselage resistance of the flight and is enclosed by the bladed rotor R. F. , which is in one of the rotor levels RP this bladed rotor R perpendicular to its geometric axis of rotation RA results.

In accordance with an important aspect of the invention, here both rotors are R, with respect to the wing assembly W. in the direction of flight FR positioned in front of this. This means that the aerodynamic flow onto the rotor R cannot pass through the wing arrangement W. be disturbed. Thus, in contrast to the prior art, there is no turbulence in the lift-generating wing W. into one of the bladed rotors R, which experience has shown would greatly increase the noise emitted by the bladed rotors R. In addition, the bladed rotors R arranged in front of the airfoil experience a largely homogeneous speed distribution in their flow over their rotor surface. By arranging the bladed rotors R in front of the wing arrangement W. a further low-noise operation of the aircraft succeeds according to the invention. Due to the arrangement of two bladed rotors R with opposite directions of rotation in front of the wing arrangement W. the flow of air onto the wing arrangement succeeds W. In contrast to the state of the art, this is also advantageously largely free of twist. Another important characteristic of the invention can be seen in this embodiment, namely that the geometric axis of rotation RA at least one bladed rotor R, viewed in a plane immediately upstream of the airfoil assembly W. and perpendicular to the axis of rotation RA this bladed rotor R lies and, along the direction of the aircraft vertical axis VA from the ground level BO seen from above the resulting minimum vertical level in this plane of the upper surface of the wing assembly W. , is arranged so that in a rotor plane RP This one bladed rotor R is the predominant area that is in circulation by the rotor blades B. of the bladed rotor R swept geometric rotor surface RAE , from the ground level BO viewed from above the wing assembly W. comes to rest in such a way that those facing in the opposite direction of flight RFD downstream wing assembly W. caused the ram pressure increase of the propeller jet at least this bladed rotor R predominantly above the wing arrangement W. he follows.

Another very important advantage according to the invention results, as shown here, from the fact that the rotors R in the direction of flight FR the wing assembly W. are connected upstream and likewise the tail assembly LW of the wing assembly W. is downstream in the direction of flow. This allows the rotor R, aided by the retractable landing gear of the landing gear arrangement GA , for example, for maintenance, repair or replacement, can be pulled forward over the fuselage, even in one piece. It no longer has to be dismantled for this. According to this embodiment, the aircraft has a tail unit arrangement, this consisting of a horizontal tail unit arrangement and a vertical tail unit arrangement. This tail arrangement is in the direction of the longitudinal axis of the fuselage FA , in the direction of flight FR seen, in front of and in the direction of flow behind the wing assembly W. so installed according to the invention that both the horizontal stabilizer arrangement and the vertical stabilizer arrangement are arranged at least in sections in the direct wake of the propeller jet generated by at least one bladed rotor R. The tail units are at least partially located in the propeller jet of a rotor, subjected to an increased dynamic pressure and are therefore more effective. In this exemplary embodiment shown, there is the motor device E. of two motors, each bladed rotor R for supplying power to at least one of these two specific motors of the motor device E. assigned. More precisely, a motor drives the motor assembly here E. over a long distance wave T and about gears G a first bladed rotor R on. A second motor of the motor assembly E. also drives via an assigned long-distance shaft T and the second bladed rotor R via gear. The distant waves T are on opposite sides of the fuselage assembly F. , partly stored on this, arranged. A special feature of this embodiment is that the transmission G is partly integrated on the rotating rotor ring. This rotor ring of the bladed rotor here has an internal toothing, in which another Gear engages, which is in kinematic operative connection with the long-distance shaft T stands by a motor of the motor assembly E. is driven. The active kinematic connection can be done directly, for example via a shaft-hub connection or generally also via further gears of the transmission G , respectively. The rotor ring thus simultaneously forms the larger gear wheel of a reduction gear between the bladed rotor R and the motor arrangement E. . The two motors of the motor assembly E. are in one of the rotor plane RP perpendicular to its axis of rotation RA a bladed rotor R considered, arranged so that at least one motor in an area between the nose landing gear leg ATM and the right main landing gear leg of the main landing gear assembly GAM and at least one second motor in an area between the nose gear leg ATM and the left one of the main landing gear legs of the main landing gear assembly GAM is arranged. In this way, the motors of the motor assembly E. , also in the direction of the wingspan of the wing arrangement W. seen, be advantageously installed close to the center of gravity. At the same time, the two motors are the motor arrangement E. in this embodiment, in the opposite direction of flight RFD , thus seen approximately in the direction of flow, along the longitudinal axis of the fuselage FA arranged behind at least one bladed rotor R formed rotor system. In this way, the motors of the motor assembly E. in the direction of the longitudinal axis of the fuselage FA advantageously be installed close to the center of gravity. In addition, the motors are the motor device E. along the vertical axis VA of the aircraft, in the vertical direction from the ground plane BO seen from below the wing assembly W. arranged. In this way, too, the motors of the motor assembly E. in the direction of the vertical axis VA of the aircraft are advantageously installed close to the center of gravity. The two motors of the motor device E. are designed in this example as a gas turbine that primarily emits shaft power. Each of the two bladed rotors R is here via at least one gear G via an associated motor of the motor assembly E. driven, these driving motors of the motor assembly E. proportionally in the fuselage arrangement F. , but outside the pressurized cabin P. , are arranged. They are in an area of the aircraft outside the pressurized cabin P. , but arranged sunk at least in part within one of the space that is exposed to the outside to the fluid through the from the outer skin of the wing assembly W. and that of the outer skin of the fuselage assembly F. formed outer contour of the aircraft is delimited, and so arranged embedded in this space that a component that in the function of an engine mount structural forces of at least one engine of the engine assembly E. in the direction of the structure of the aircraft, is essentially not visible from the outside and that this component in its function of the engine mount outside of the by the outer skin of the wing assembly W. and from the skin of the fuselage assembly F. formed outer contour, essentially no separate, resistance-effective surface flushed by the fluid can be assigned. In this embodiment, the shaft power turbines in the vicinity of the wing / fuselage area, that is to say in the area of the belly fairing, are partially sunk in an unprinted area U arranged, which emerges here, for example, from an expansion of that unprinted space in which the landing gear shaft is also arranged. The air inlets are suitable far from the unprinted area U and the engine exhaust gases are also discharged from this space via guides. Thus, these motors cause the motor assembly E. , now mostly inside the aircraft, significantly less air resistance, which also reduces the aircraft's fuel consumption again. The shaft drives attached to drive the bladed rotors are thus part of the motor arrangement E. sunk and mounted very close to the center of gravity with regard to all coordinate axes. This in turn means for the aircraft, also with the rotors R, which are also attached here close to the center of gravity, that lower moments of inertia result for the aircraft. In addition, the tail unit arrangement is designed as a conventional cross tail unit. The external noise of the propeller turbine is greatly reduced by the recessed arrangement in the aircraft. According to one embodiment of the invention, the shaft power turbines that drive the bladed rotors generate additional thrust in the sense of a propulsive force for the aircraft. The exhaust gas jet (indicated here with a gray dashed arrow) can also be led out in this exemplary embodiment on the underside of the fuselage rear in such a way that it can energize the boundary layer in the area of the rear fuselage constriction so energetically that an aerodynamic separation is reduced, which is the Resistance of the trunk with could decrease. According to an inventive feature, both bladed rotors R are between a nose gear assembly ATM and a main landing gear assembly GAM in the direction of the longitudinal axis of the fuselage FA on the fuselage assembly F. arranged so that in each of the rotor planes RP considering each of these bladed rotors R, the largest outside diameter OD of the bladed rotor R is smaller than the one with the undercarriage arrangement GA specified, double the minimum distance of the axis of rotation RA this bladed rotor R from the ground plane BO . In this way, the bladed rotor R can safely rotate in all of the usual phases of operation flush with the ground, without being flush with the ground BO to get in touch. According to this embodiment, the largest outside diameter is OD of the bladed rotors selected so that the bladed rotor R also has a safety distance of around 60 cm from the ground level BO having.

2 The figure shows, in a greatly simplified form, an exemplary embodiment of a future commercial aircraft according to the invention in a front view. According to an important aspect of the invention, two bladed rotors R are a fuselage arrangement that can be printed at least in sections F. , this annularly surrounding rotatably arranged that each rotor R in one of its rotor planes RP , which in the area of its length along its rotor axis RA are arranged perpendicular to this, at least part of a pressure cabin P. that is inside the fuselage arrangement F. lies, surrounds. According to the invention, this has the consequence that in the event of a rotor breakage during operation, fragments of the rotor R move away from the fuselage arrangement F. , but especially the pressurized cabin P. , remove by the effective centrifugal forces. According to the invention, each rotor surrounds R 1 in one of its rotor planes RP seen, here that cross section of the fuselage assembly F. , which in this plane at least in sections according to a pressure cabin P. is printable. The bladed rotors R are rotatable in relation to the fuselage arrangement F. by at least one bearing arrangement S. (as shown in ) stored and approximately in the direction of the longitudinal axis of the fuselage FA by at least one bearing arrangement S. (as shown in ) fixed arranged. In this exemplary embodiment, the axes of rotation are RA , which coincidentally coincide here with these two rotors R, in the sense of a usual drive camber at a small angle in relation to the longitudinal axis of the fuselage FA inclined to run. The two rotors R are inventive along the longitudinal axis of the fuselage FA on the fuselage assembly F. arranged that in all of the rotor planes RP of each of the two rotors that are perpendicular to the axis of rotation RA of the bladed rotor R and are arranged within its longitudinal extent, no further important safety-critical components necessary for the immediately safe flight of the aircraft are arranged. Thus, in the event of a bladed rotor R breaking, these cannot be damaged in a safety-critical manner. Also here, according to a special characteristic of the invention, both bladed rotors R are designed in such a way that they rotate freely in the fluid without a jacket. As a result, they achieve a very high level of efficiency in the propulsion conversion. In this application example, the two bladed rotors R are advantageously installed in such a way that they rotate in opposite directions to one another. Together they form a rotor system. In this way the rotor efficiency of the rotor system is improved. The bladed rotors R are on the fuselage assembly F. , these not only arranged in a surrounding way, but in a certain way compared to a 3-point chassis arrangement GA , with a total of several rotatable wheels WE and with a nose landing gear ATM , in which GA with the statically determined support of the aircraft against the ground level BO serves in operation flush with the ground, positioned so that both bladed rotors R on the fuselage assembly along the longitudinal axis of the fuselage FA between a nose landing gear structurally attached to the fuselage assembly ATM and a main landing gear assembly GAM are positioned, the latter being attached to the wing assembly W. is structurally connected and includes two main landing gear legs. Since the rotation during take-off and the derotation during landing are essentially centered around the wheels of the main landing gear assembly GAM takes place, there is therefore no risk according to the invention during these critical phases that a bladed rotor R on the ground level BO touches down or its ground clearance is impermissibly reduced during the rotation or derotation of the aircraft. Corresponding to this design, the dimension for the generation of propulsion is that of the rotor blades B. of a bladed rotor R in rotation swept over the geometric rotor surface RAE designed to be the same size or larger than the cross-sectional area of the fuselage arrangement that is considered a measure of the fuselage resistance of the flight and is enclosed by the bladed rotor R. F. , which is in one of the rotor levels RP this bladed rotor R perpendicular to its geometric axis of rotation RA results.

In accordance with an important aspect of the invention, here both rotors are R, with respect to the wing assembly W. in the direction of flight FR positioned in front of this. This means that the aerodynamic flow onto the rotor R cannot pass through the wing arrangement W. be disturbed. Thus, in contrast to the prior art, there is no turbulence in the lift-generating wing W. into one of the bladed rotors R, which experience has shown would greatly increase the noise emitted by the bladed rotors R. In addition, the bladed rotors R arranged in front of the airfoil experience a largely homogeneous speed distribution in their flow over their rotor surface. By arranging the bladed rotors R in front of the wing arrangement W. a further low-noise operation of the aircraft succeeds according to the invention. Due to the arrangement of two bladed rotors R with opposite directions of rotation in front of the wing arrangement W. the flow of air onto the wing arrangement succeeds W. In contrast to the state of the art, this is also advantageously largely free of twist. Another important characteristic of the invention can be seen in this embodiment, namely that the geometric axis of rotation RA at least one bladed rotor R, viewed in a plane immediately upstream of the airfoil assembly W. and perpendicular to the axis of rotation RA this bladed rotor R lies, and along the direction of the aircraft vertical axis VA from the ground level BO seen from above the resulting minimum vertical level in this plane of the upper surface of the wing assembly W. , is arranged so that in a rotor plane RP This one bladed rotor R is the predominant area that is in circulation from the rotor blades B. of the bladed rotor R swept geometric rotor surface RAE , from the ground level BO viewed from above the wing assembly W. comes to rest in such a way that those facing in the opposite direction of flight RFD downstream wing assembly W. caused ram pressure increase of the propeller jet at least this bladed rotor R predominantly above the wing arrangement W. he follows.

Another very important advantage according to the invention results, as shown here, from the fact that the rotors R in the direction of flight FR the wing assembly W. are connected upstream and likewise the tail assembly LW of the wing assembly W. is downstream in the direction of flow. This allows the rotor R, aided by the retractable landing gear of the landing gear arrangement GA , for example, for maintenance, repair or replacement, can be pulled forward over the fuselage, even in one piece. It no longer has to be dismantled for this. According to this embodiment, the aircraft has a tail unit arrangement, this consisting of a horizontal tail unit arrangement and a vertical tail unit arrangement. This tail arrangement is in the direction of the longitudinal axis of the fuselage FA , in the direction of flight FR seen, in front of and in the direction of flow behind the wing assembly W. so installed according to the invention that both the horizontal stabilizer arrangement and the vertical stabilizer arrangement are arranged, at least in sections, in the direct wake of the propeller jet generated by at least one bladed rotor R. The tail units, at least partially located in the propeller jet of a rotor, are subjected to an increased dynamic pressure and are therefore more effective. In this exemplary embodiment shown, there is the motor device E. of two motors, each bladed rotor R for supplying power to at least one of these two specific motors of the motor device E. assigned. More precisely, a motor drives the motor assembly here E. over a long distance wave T and about gears G a first bladed rotor R on. A second motor of the motor assembly E. also drives via an assigned long-distance shaft T and the second bladed rotor R via gear. The distant waves T are on opposite sides of the fuselage assembly F. , partly stored on this, arranged. A special feature of this embodiment is that the transmission G is partly integrated on the rotating rotor ring. This rotor ring of the bladed rotor here has an internal toothing in that a further gear wheel engages, which is in kinematic operative connection with the remote shaft T stands by a motor of the motor assembly E. is driven. The active kinematic connection can be done directly, for example via a shaft-hub connection or generally also via further gears of the transmission G , respectively. The rotor ring thus simultaneously forms the larger gear wheel of a reduction gear between the bladed rotor R and the motor arrangement E. . The two motors of the motor assembly E. are in one of the rotor plane RP perpendicular to its axis of rotation RA a bladed rotor R considered, arranged so that at least one motor in an area between the nose landing gear leg ATM and the right main landing gear leg of the main landing gear assembly GAM and at least one second motor in an area between the nose gear leg ATM and the left one of the main landing gear legs of the main landing gear assembly GAM is arranged. In this way, the motors of the motor assembly E. also, in the spanwise direction of the wing arrangement W. seen, be advantageously installed close to the center of gravity. At the same time, the two motors are the motor arrangement E. in this embodiment, in the opposite direction of flight RFD , thus seen approximately in the direction of flow, along the longitudinal axis of the fuselage FA arranged behind at least one bladed rotor R formed rotor system. In this way, the motors of the motor assembly E. in the direction of the longitudinal axis of the fuselage FA advantageously be installed close to the center of gravity. In addition, the motors are the motor device E. along the vertical axis VA of the aircraft, in the vertical direction from the ground plane BO seen from below the wing assembly W. arranged. In this way, too, the motors of the motor assembly E. in the direction of the vertical axis VA of the aircraft are advantageously installed close to the center of gravity. The two motors of the motor device E. are designed in this example as a gas turbine that primarily emits shaft power. Each of the two bladed rotors R is here via at least one gear G via an associated motor of the motor assembly E. driven, these driving motors of the motor assembly E. proportionally in the fuselage arrangement F. , but outside the pressurized cabin P. , are arranged. They are in an area of the aircraft outside the pressurized cabin P. , but at least partially sunk within one of the space that leads to the outside to the fluid through the from the outer skin of the wing assembly W. and that of the outer skin of the fuselage assembly F. formed outer contour of the aircraft is delimited and so arranged embedded in this space that a component that in the function of an engine mount structural forces of at least one engine of the engine assembly E. in the direction of the structure of the aircraft, is essentially not visible from the outside and that this component in its function of the engine mount outside the, through the from the outer skin of the wing assembly W. and from the skin of the fuselage assembly F. formed outer contour essentially no separate, resistance-effective surface flushed by the fluid can be assigned. In this embodiment, the shaft power turbines in the vicinity of the wing / fuselage area, that is to say in the area of the belly fairing, are partially sunk in an unprinted area U arranged, which emerges here, for example, from an extension of the unprinted space in which the landing gear shaft is also arranged. The air inlets are suitable far from the unprinted area U and the engine exhaust gases are also discharged from this space via guides. Thus, these motors cause the motor assembly E. , now mostly inside the aircraft, significantly less air resistance, which also reduces the aircraft's fuel consumption again. The shaft drives attached to drive the bladed rotors are thus part of the motor arrangement E. , sunk and mounted very close to the center of gravity with regard to all coordinate axes. This in turn means for the aircraft, also with the rotors R, which are also attached here close to the center of gravity, that lower moments of inertia result for the aircraft. In addition, the tail unit arrangement is designed as a conventional cross tail unit. The external noise of the propeller turbine is greatly reduced by the recessed arrangement in the aircraft. According to one embodiment of the invention, the shaft power turbines that drive the bladed rotors generate additional thrust in the sense of a propulsive force for the aircraft. The exhaust gas jet (indicated here with a gray dashed arrow) can also be led out in this exemplary embodiment on the underside of the fuselage tail in such a way that it can energetically revive the boundary layer in the area of the rear fuselage constriction so that aerodynamic detachment is reduced, which is the Resistance of the trunk with could decrease. According to an inventive feature, both bladed rotors R are between a nose gear assembly ATM and a main landing gear assembly GAM in the direction of the longitudinal axis of the fuselage FA on the fuselage assembly F. is arranged so that in each of the rotor planes RP , considering each of these bladed rotors R, the largest outside diameter OD of the bladed rotor R is smaller than the one with the undercarriage arrangement GA specified, double the minimum distance of the axis of rotation RA this bladed rotor R from the ground plane BO . In this way, the bladed rotor R can safely rotate in all of the usual phases of operation flush with the ground, without being flush with the ground BO to get in touch. According to this embodiment, the largest outside diameter is OD of the bladed rotors selected so that the bladed rotor R also has a safety distance of around 60 cm from the ground level BO having.

3 an overview of the previously common printed areas P. and unprinted areas U within the fuselage assembly F. a transport aircraft; being those in the fuselage assembly F. lying parts, which are shown in white, as a pressure cabin P. can be printed, whereby, according to the invention, further unprinted areas can be created or existing ones can be enlarged.

4th as in the brief description above

5 a greatly simplified, exemplary rotatable arrangement of a rotor R according to the invention with a bearing arrangement S. , the profile section of the fuselage arrangement F. surrounding ring-shaped, being within the cross-section of the fuselage assembly F. one according to a pressurized cabin P. printable section is provided which depends on its vertical height V is designed so high that an average passenger can stand upright in it; whereby only the upper part of the fuselage cross-section is printed here according to a pressure cabin, while the lower part is separated from the upper part by a kind of floor arrangement. The bladed rotor has a smallest inner diameter ID and a largest outside diameter OD on.

6th Profile section of an exemplary and greatly simplified embodiment of a bearing and guide unit with a bearing arrangement S. for the rotary mounting of the drive rotor R with simultaneous drive-effective dissipation of the axial forces and kinematic coupling of the bladed rotor R via a gearbox to a remote shaft T , with the wave power over the long-distance wave T Via bevel gears in an internally toothed ring according to a gear G engages with the internally toothed ring being part of the fuselage assembly F. orbiting bladed rotor R is;

7th a greatly simplified exemplary kinematic coupling of a bladed rotor R according to the invention via gears to several motors of the motor arrangement E. , wherein the shoved rotor R has an internal toothing on which a further gear, which is connected to a motor of the motor arrangement E. is connected, acts kinematically, here by way of example as a hybrid-electric embodiment of the drive system, with a motor of the motor arrangement E. is designed to be engaged and disengaged via a clutch; one motor being designed as a gas turbine and the further motor being designed as an electric motor, with only a few of all rotor blades in a simplified representation B. are shown on the rotor;

8th shows a greatly simplified exemplary sketched representation of the possible arrangement of a bladed rotor, shown in a plane at a fuselage length position at the level of the wing arrangement W. downstream of the main landing gear assembly GAM perpendicular to the axis of rotation of a bladed rotor, here in the direction of flight FR seen, the largest outside diameter OD at least one bladed rotor is designed to be equal to or smaller than twice the maximum longitudinal extension LGAM a landing gear leg with wheels WE the main landing gear arrangement GAM . This is an arrangement of the main landing gear GAM possible in such a way that the two landing gear legs attached to the wings neither interfere with the inflow nor the outflow of a rotor R. Likewise, as indicated here, the landing gear can be pivoted inward and retracted. Due to this geometric arrangement, the available space is also sufficient in length for the retraction of both landing gear legs. At the same time, the rotor cannot geometrically touch the ground plane. In this figure, it is also shown in this plane as an example that the axis of rotation is in this plane RA of a rotor along the aircraft vertical axis VA , from the ground level BO seen above a vertical minimum MV the top of the wing assembly W. comes to rest so that the predominant part (indicated here by dashed lines) of the rotor blades also shown here B. area swept over the entire circumference RAE above the wing assembly W. comes to rest. The distance between the landing gear legs can of course also be made larger than shown here, but should at the same time also be larger than the largest outer diameter of the bladed rotor R.

9 shows a greatly simplified exemplary sketched representation of a with rotor blades B. bladed rotor R, on the left in the side view and on the right in the front view and illustrates the definition of rotor planes RP in the area of the longitudinal extent of the bladed rotor R along its axis of rotation RA , these rotor planes RP always to the axis of rotation RA stand normally. As shown in the illustration on the left, a first rotor plane RP normal to the axis of rotation RA of the bladed rotor R can be defined in front, which is in the direction of flight FR seen, while just touching the foremost tip of a foremost rotor blade of the rotor R, a last rotor plane RP with regard to the longitudinal extension of the rotor R along its axis of rotation RA touches, just seen in the direction of flow, one end of a rotor blade that is positioned as far back as possible in the direction of flow B. and this plane is of course also perpendicular to the axis of rotation RA of the bladed rotor R. Between these two planes, front and rear planes, any number of planes can be created within the longitudinal extent of the rotor along its axis of rotation RA are formed (the corresponding area in which planes of rotation RP can occur at all is indicated by arrows between the two levels). Each of these planes is normal to the axis of rotation RA and reaches an extent in the area which exceeds the spatial extent of the aircraft. The planes of rotation RP serve as geometric auxiliary planes to determine the possible area of impact of a bladed rotor R on the aircraft.

Claims (21)

Aircraft with: - an at least partially cylindrically elongated fuselage arrangement (F) with a spatial main extent in the direction of a longitudinal fuselage axis (FA), this fuselage arrangement (F) for accommodating payload as required, the fuselage arrangement (F) at least in sections functioning as a pressurized cabin ( P) can be imprinted with an internal pressure, - a wing assembly (W) firmly attached to the fuselage assembly (F) of the aircraft, for generating lift supporting the aircraft, the force component of the lift generated in horizontal flight perpendicular to the longitudinal axis of the fuselage (FA) opposite the component along the longitudinal axis of the fuselage (FA) clearly predominates, - with a three-point landing gear arrangement (GA) that can be retracted and extended again if necessary, with several rotatably mounted wheels (WE) for statically determined support of the aircraft against the ground level (BO) in floor-level operation, consisting of two structurally the wing assembly (W) connected main landing gear legs (GAM) and a nose landing gear leg (GAA) structurally connected to the fuselage arrangement (F), the nose landing gear leg (GAA) being arranged along the longitudinal axis (FA) in flight direction (FR) in front of the main landing gear legs (GAM), a motor arrangement (E), consisting of at least one motor, for generating a drive power available to the aircraft, - at least one rotor (R) bladed with several rotor blades (B) for at least partial conversion of drive power of the motor arrangement (E) into a Propulsion power of the aircraft through rotation of the bladed rotor (R) around a geometric axis of rotation (RA), with a geometric rotor surface (RAE) swept over by the rotor blades (B) when the rotor blades (B) of the bladed rotor (R) rotate. results, this bladed rotor (R) in its radial spatial extension a smallest inner diameter (ID) and e having the largest outer diameter (OD), this bladed rotor (R) being equipped with blades (B) which are adjustable in the setting angle, this bladed rotor (R) also having a spatial longitudinal extension along its geometric axis of rotation (RA), within this range its longitudinal extension Rotor planes (RP) perpendicular to its geometric axis of rotation (RA) can be identified as a geometric aid to detect a possible area of impact of this rotor (R), - at least one gear (G) for transmitting the drive power of at least one motor of the motor arrangement (E) At least one bladed rotor (R), - at least one bearing arrangement (S), characterized in that at least one bladed rotor (R) with at least one bearing arrangement (S) is rotatably mounted surrounding the fuselage arrangement (F) and that at least this one Bladed rotor (R) is axially fixed approximately in the direction of the fuselage longitudinal axis (FA) by at least one bearing arrangement (S) to the fuselage arrangement (F), and that at least this one bladed rotor (R) at the same time in the radial direction outside that section of one at least partially Cylindrical elongated fuselage arrangement (F) of an aircraft surrounds ring-shaped, which, in one of the red or planes (RP) of this rotor (R) considered, can be printed at least partially in the function of a pressure cabin (P) and that at least in one of the rotor planes (RP) of the bladed rotor (R) no other essentials for the immediately safe operation of the aircraft Components of the aircraft are arranged, and that at least this one bladed rotor (R) along the fuselage longitudinal axis (FA), seen in the direction of flight (FR), is arranged in front of the wing arrangement (W) and at the same time the geometric axis of rotation (RA) of at least one bladed rotor (R), viewed along a direction of the aircraft vertical axis (VA), starting from the ground plane (BO), viewed in a plane that is immediately upstream of the wing arrangement (W) and perpendicular to the axis of rotation (RA) of this bladed rotor (R) , is arranged so that it is above the minimum vertical level (MV) of the upper surface of the wing assembly (W) resulting in this plane, b e finds that in a rotor plane (RP) of this one bladed rotor (R), the predominant area portion of the geometric rotor area (RAE) swept over by the rotor blades (B) of the bladed rotor (R), from the bottom level (BO) considered, above the wing arrangement (W) comes to rest in such a way that the ram pressure increase of the propeller jet caused on the wing arrangement (W) arranged downstream against the direction of flight (FR) at least this one bladed rotor (R) occurs predominantly above the wing arrangement (W) and that at least this one rotor (R) in the direction of the fuselage longitudinal axis (FA) on the fuselage arrangement (F) between the nose landing gear leg (GAA) and the main landing gear legs of the main landing gear assembly (GAM) is arranged in such a way that this is in each of the rotor levels (RP) Considering a bladed rotor (R), the largest outside diameter (OD) of this one rotor (R) is smaller than that with the undercarriage arrangement (GA) vo The given, double the minimum distance of the axis of rotation (RA) of this bladed rotor (R) from the ground plane (BO), and that at least one motor of the motor device (E) is in an area of the aircraft outside the pressure cabin (P), but at least partially is arranged sunk within a space which is delimited to the outside from the fluid by the outer contour of the aircraft formed by the outer skin of the wing assembly (W) and the outer contour of the aircraft formed by the outer skin of the fuselage assembly (F), and is arranged so embedded in this space that a Component which, in the function of an engine bracket, dissipates structural forces from at least one engine of the engine assembly (E) in the direction of the structure of the aircraft, is essentially not visible from the outside and that this component in its function of the engine bracket outside of the by the Outer skin of the wing arrangement (W) and the outer contour formed by the outer skin of the fuselage arrangement (F) essentially no e Its own, resistive surface flushed by the fluid can be assigned, and that at least this one bladed rotor (R) is kinematically coupled to at least one motor of the motor arrangement (E), which is arranged outside the pressure cabin (P), via a gear (F) , and can be set in rotation by this in order to generate a propulsive force that drives the aircraft, the force component of which in the direction of the longitudinal axis of the fuselage (FA) clearly outweighs the force component perpendicular to the longitudinal axis of the fuselage (FA). Plane to Claim 1 with a tail assembly (LW), which is spaced in the direction of the longitudinal axis of the fuselage (FA) from the wing assembly (W) and is in the opposite flight direction (FR) behind the wing assembly (W), consisting of a horizontal stabilizer assembly and a vertical stabilizer assembly, whereby both the horizontal stabilizer assembly and the Vertical tail arrangement are arranged at least in sections in the direct wake of the propeller jet generated by at least one bladed rotor (R). Plane to at least one of the Claims 1 - 2 , whereby the vertical height (V) of the pressure cabin (P) within the fuselage arrangement (F) is designed in such a way that an upright gait is possible there for the average passenger, at least in sections, and the smallest inner diameter (ID) of at least one bladed rotor ( R) is at least larger in terms of its dimensions than this vertical construction height (V) of the pressure cabin (P). Plane to at least one of the Claims 1 - 3 , at least one coupling in the kinematic drive train between at least one bladed rotor (R) and at least one motor of the motor arrangement (E), which enables the bladed rotor (R) to be decoupled from at least one motor of the motor device (E) if necessary. Plane to at least one of the Claims 1 - 4th wherein the motor device (E) consists of a plurality of motors, each bladed rotor (R) being assigned at least one specific motor of the motor device (E) for supplying power. Plane to at least one of the Claims 1 - 5 wherein at least one motor of the motor arrangement (E), seen in the opposite direction of flight (RFD), is arranged along the longitudinal axis of the fuselage behind a rotor system formed from at least one bladed rotor (R). Plane to at least one of the Claims 1 - 6th wherein at least one motor of the motor device (E) is arranged below the wing arrangement (W) along the vertical axis (VA) of the aircraft, viewed in the vertical direction from the ground plane (BO). Plane to at least one of the Claims 1 - 7th , wherein at least one motor of the motor device (E) is designed as a gas turbine that delivers primarily shaft power. Plane to at least one of the Claims 1 - 8th At least one motor of the motor arrangement (E) is designed as a gas turbine, the thermal efficiency of which is further increased in the sense of unconventional measures by recuperative devices for intermediate compressor cooling, fuel mass preheating or exhaust gas heat exchange. Plane to at least one of the Claims 1 - 9 , wherein the motor arrangement (E) consists of several motors, with a rotor plane (RP) perpendicular to its axis of rotation (RA) of a bladed rotor (R) viewed, these motors are arranged so that at least one motor in an area between the nose landing gear leg (GAA) and the right of the main landing gear (GAM) and at least one second motor is arranged in an area between the nose landing gear (GAA) and the left of the main landing gear (GAM). Plane to at least one of the Claims 1 - 10 , wherein at least one bladed rotor (R) rotates freely in the fluid without a jacket. Plane to at least one of the Claims 1 - 11 , an even number of several bladed rotors being arranged on the fuselage arrangement (F), two of these bladed rotors (R) each having opposite directions of rotation during operation. Plane to at least one of the Claims 1 - 12th , whereby the geometric rotor area (RAE) covered in circulation by the rotor blades (B) of a bladed rotor (R), which is the measure for the generation of propulsion, is the same size or larger than that which is used as the measure for the fuselage resistance of the flight and the bladed Rotor (R) enclosed cross-sectional area of the fuselage arrangement (F), which results in one of the rotor planes (RP) of this bladed rotor (R) perpendicular to its geometric axis of rotation (RA). Plane to at least one of the Claims 1 - 13th , whereby at least one geometric rotor area (RAE), which is valid as a measure for the generation of propulsion and which is swept over by the rotor blades (B) of a bladed rotor (R) in circulation, is designed to be the same size or larger than fifteen times the sum of the air inlet cross-sectional areas of all air-breathing engines of the Motor arrangement (E) that drive this bladed rotor (R) and that the rotor blades (B) of this bladed rotor (R) are designed aerodynamically so that the aircraft can reach flight speeds of at least Ma 0.78 at altitudes greater than 10,000 m. Plane to at least one of the Claims 1 - 14th At least one bladed stator is connected upstream or downstream of a bladed rotor (R), approximately in the direction of the aircraft longitudinal axis (FA) and is designed to convert the swirl energy of the air flow at least partially into propulsion for the aircraft. Plane to at least one of the Claims 1 - 15th , whereby the individual rotor blades (B) of at least one bladed rotor (R) can be adjusted independently of one another in their setting angle to the air flow, and this can also vary over the circumferential position of the rotor so that a vectorization of the thrust results, which is also used for the control , for trimming, to save drag and to reduce the noise of the aircraft. Plane to at least one of the Claims 1 - 16 , wherein at least one bladed rotor (R) is driven by at least one ring-shaped electric motor, which is arranged approximately concentrically to the bladed rotor (R), also ring-shaped surrounding the fuselage arrangement (F). Plane to at least one of the Claims 1 - 17th , whereby the nose landing gear assembly (GAA) can be retracted and extended independently of the main landing gear assembly (GAM). Plane to at least one of the Claims 1 - 18th , the largest outer diameter (OD) of at least one bladed rotor (R) being the same size or smaller than twice the maximum longitudinal extension of a landing gear leg including the wheels (WE) of the main landing gear assembly (GAM). Plane to at least one of the Claims 1 - 19th , wherein at least one motor of the motor arrangement (E), which drives at least one bladed rotor (R), is designed as an air-breathing motor and at least this generates further propulsive force in the form of thrust through the ejection of exhaust gases in addition to the propulsive force of at least one rotor (R) With this propulsion force, too, the force component in the longitudinal direction of the axis of rotation (RA) clearly outweighs the force component perpendicular to the axis of rotation (RA). Use of an aircraft according to at least one of the Claims 1 - 20th .

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