WO2014108125A1 - Aéronef à rendement élevé et à faible bruit - Google Patents

Aéronef à rendement élevé et à faible bruit Download PDF

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Publication number
WO2014108125A1
WO2014108125A1 PCT/DE2014/000007 DE2014000007W WO2014108125A1 WO 2014108125 A1 WO2014108125 A1 WO 2014108125A1 DE 2014000007 W DE2014000007 W DE 2014000007W WO 2014108125 A1 WO2014108125 A1 WO 2014108125A1
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WO
WIPO (PCT)
Prior art keywords
rotor
assembly
aircraft
bladed
fuselage
Prior art date
Application number
PCT/DE2014/000007
Other languages
German (de)
English (en)
Inventor
Malte Schwarze
Thomas ZÖLD
Original Assignee
Malte Schwarze
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Priority claimed from DE102013015364.6A external-priority patent/DE102013015364A1/de
Application filed by Malte Schwarze filed Critical Malte Schwarze
Priority to DE112014000391.3T priority Critical patent/DE112014000391B4/de
Publication of WO2014108125A1 publication Critical patent/WO2014108125A1/fr

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D35/00—Transmitting power from power plants to propellers or rotors; Arrangements of transmissions
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C1/00—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like
    • B64C1/16—Fuselages; Constructional features common to fuselages, wings, stabilising surfaces or the like specially adapted for mounting power plant
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
    • B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
    • B64D27/02—Aircraft characterised by the type or position of power plants

Definitions

  • the invention relates to an aircraft, preferably an aircraft of the type of commercial aircraft.
  • This printable section is printed for cruise as a pressurized cabin and allows the aircraft, with crew as well as with passengers or cargo, to seek greater altitudes in order to operate there under more favorable aerodynamic and economic conditions.
  • the overall height of the pressure cabin within the fuselage assembly is often such that for practical and comfort reasons within the pressure cabin, at least in sections, an upright gear is possible for the average passenger.
  • a key objective of some stakeholders is to further increase the fuel efficiency of transport aircraft, in particular commercial aircraft, and reduce their noise emitted during operation.
  • turbofan engines allow the side stream to move much of the air mass flow through the engine at a comparatively low flow rate through a fan to guide it around the engine core.
  • the fan which contributes to the acceleration of the majority of the mass flow and thus also to a large part of the thrust generation, is thereby driven by the core turbine in the main flow, which removes the mechanical work from the fluid and mechanically supplies it to the fan via a shaft.
  • the drive-generating fan in the sidestream are functionally directly coupled and integrated in close proximity to one another within a single engine housing.
  • thrust is physically defined as a force as a change in momentum.
  • this change in the momentum and thus the generation of a certain thrust on the one hand can be done, in principle, by the acceleration of a relatively small mass flow with a high change in speed. This possibility causes in the outlet stream a high jet velocity and thus a comparatively high jet noise.
  • This type of propulsion generation is especially suitable for achieving high airspeeds, such as combat jets.
  • the same push level can also be achieved by giving a fairly 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.
  • This thrust cross-sectional area of the propulsion-inducing fan may also be called a rotor surface and, in many cases, such as a propeller, also corresponds approximately to the inlet cross-sectional area which a mass flow passes through at the engine to participate in propulsion generation. This is simplistic and clearly imaginable.
  • the mass flow, which is to be increased here, is defined as the mass of air which flows through a certain thrust cross-sectional area or rotor surface 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 for thrust generation. If the density of the fluid and the speed of flight (which roughly corresponds to the speed of entry into the rotor surface) remain approximately the same in a first approach, then the geometric rotor area must be further increased in order to achieve an unchanged thrust level. However, an increase in the rotor area is only possible by increasing the diameter of the drive-generating rotor.
  • the fan is decoupled from the rotational speed of the drive and turbine shaft by means of a reduction gearbox. Due to the fan's lower spin speed, the fan's blade tip number is reduced. and increase its component efficiency.
  • the engine can be designed for a lower fan pressure ratio, whereby the fluid in the bypass flow is accelerated less and flows slower. 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.
  • the shear cross-sectional area, here the thrust-generating rotor surface of the fan must be increased. Due to the lower rotational angular velocity of the fan, however, its diameter and thus its thrust area can be increased without reaching the blade tips criticaltown.ergeschwihdtechniken. Thus, a low-noise operation of the fan is possible, with the fan noise results approximately from the fourth power of the rotational speed of its leaves.
  • the improvement in specific fuel consumption is achieved mainly by improving the propulsive efficiency as a result of the increased by-pass ratio. Disadvantage is that nevertheless the rotor surface of the propulsion-generating fans must be increased and thus the engine nacelle fails larger and more resistant to resistance.
  • Open rotor configurations have the potential to significantly increase the propulsive efficiency and thus achieve significantly higher bypass ratios and significantly higher fuel efficiency.
  • the rotor that produces the impeller rotates freely and is no longer surrounded by a sheathing casing. This initially eliminates the resistance-generating engine nacelle.
  • an engine nacelle is necessary, which, however, is significantly smaller, approximately the same size as the diameter of fairings of previous single-stream engines, and thus significantly less resistance than conventional engine nacelles and engine gondolas from Getriebefantriebtechniken.
  • open open rotors do not contribute to an increased resistance due to their engine sheathing.
  • the O ' pen rotor technology was tested in the past at the time of the oil crises in the 70's, in the beginning already under the former term Propfan technology on flying demonstrators on the short distance in the practice, whereby at that time a fuel saving of over 20 % compared to the engines of the time.
  • Proppfan hereby emphasizes that the rotor used for propulsion is a combination of propeller and fan, combining the high propulsion efficiency of a propeller with the ability of the fan to operate at high cruising speeds. This has the consequence that proven high Reisefluggesch windtechniken 'to Ma 0.86 can be flown at a high fuel efficiency with open rotors.
  • Rotors whose blades are designed aerodynamically similar to a Propfahs, often have a low relative thickness and a continuous or progressive sweep towards the blade tip out to reach the critical Mach number only at the highest possible Auflagachiere at the blade tip. In addition, they may have an escalation for the same reason.
  • profiled sections often use advanced, transonic to supercritical pre-filings, also using the area rule.
  • the critical Machzahi is achieved at the blade tips only at relatively high rotational speed of the leaves, which means that at relatively high cruising speed, respectively Blattanströmung, up to a Mach number of up to 0.86 in a cruising altitude of over 10,000 m can be flown. This airspeed is not possible with turboprop arrangements to this extent.
  • the blades of propfan rotors are often characterized by an overall blade length that has been changed over the blade extension and increased overall.
  • the common feature of the propfan or open rotor technology that for propelling at least one open spinning rotor without sheath is used, the achieving high bypass ratios of around or above '20 is: allowing 1 and a significantly improved propulsive efficiency, often in combination with an enlarged rotor area and with the result of a significantly improved fuel efficiency.
  • Two opposing rotors also have the potential to move a large mass flow through the engine while maintaining the same fan diameter and low speed.
  • the lower Klandruck petition less rotor blades in each stage necessary what can have a favorable effect on the engine mass and on the noise, in particular with regard to the interference caused by noise between the two rotors.
  • the arrangement of open-rotor engines on the passenger plane has long been unsatisfactorily solved in total and greatly limited in their arrangement of the engines on the aircraft.
  • the vertical stabilizer is placed shielding to protect between the two rotors. In the event of damage, it should thus prevent the second rotor from being damaged after breakage in one rotor, as otherwise an engine redundancy concept would have rendered ineffective. It is controversial in the art whether the vertical stabilizer could ever be designed structurally so that it could ensure effective shielding of the rotors in practice in the event of failure.
  • the rear arrangement of the engines shown so far is very unsatisfactory overall, because not only the position of the wing assembly, but essentially the entire aircraft configuration on the engine position is determined by gravity. Due to the rear engine position, the influence on the center of gravity is great, so the loading of the aircraft with payload and fuel can not be done as flexibly as it would be desirable reduces the operational flexibility of the aircraft.
  • the tail lever arms are small, so the tail surfaces must be made large, which increases resistance and weight.
  • the moments of inertia about the transverse axis of the aircraft are also large, which further increases the tail surfaces.
  • T-tail has to be installed, which is even more difficult if it is to allow a structural shielding between the two open rotors, or if a Trimmable Horizontal Stabilizer is to be installed, which is standard on commercial aircraft.
  • the trim resistance can also be increased in cruising flight, due to the configuration of the center of gravity position and the control arm levers.
  • the open rotors have to be positioned along a longitudinal section of the aircraft along the fuselage longitudinal axis where there is no pressurized cabin, otherwise the pressurized cabin could break through fragments and be damaged in the event of a break.
  • the hitherto unsolved main problem is therefore to find the aircraft a safe arrangement form for open rotors so that essential components of the aircraft for safe flight can not be damaged.
  • it is still unsolved to find a form of arrangement for open rotors so as to maintain or improve the performance and versatility of the nuclear aircraft apart from the improved performance by the engine.
  • a propeller is installed in the fuselage and several in the sense of rotation to each other in opposite directions running propeller can be used on the fuselage.
  • the propellers are positioned in the direction of flight behind the wing assembly, so that the arrangement of weapons on the aircraft easier to maneuver and these must not be shot through the "propeller circle.”
  • the aircraft has no pressure cabin inside the fuselage
  • the engine assembly is mounted within a fuselage and drives Via a pinion, which rotates for the most part within the hull limiter, a larger-diameter circulating ring is mounted on a bearing ring, the drive being achieved by means of an internally toothed toothed wheel mounted on this raceway Since the one bearing ring of smaller diameter is rigidly connected to the hull and this over closed the full Wjnkellie, that is not running interrupted, in this arrangement, the motor pinion along the fuselage longitudinal axis in front of or behind the bearing ring, the fuselage skin must break through an opening to larger with the internally toothed ring Dia
  • tail has to be executed inappropriately small at the rear, so that the propeller can be dismantled to the rear. This has the additional consequence that the tail units are no longer illuminated by the propeller jet.
  • sternpropellers in Heckiage it is a disadvantage of sternpropellers in Heckiage that they are badly removable from the aircraft due to the tail units at the rear behind the propeller and through the front wing, and have to be disassembled for this purpose.
  • DE 1884 174 describes a power wing aircraft which applies both the lift to carry the aircraft and the propulsion to propel the aircraft over rotating rotors of great radial extent which can be kinematically pivoted with parts of the fuselage and parts thereof Hull surrounded.
  • the rotors are driven by associated blade tip drives.
  • the aircraft can start and land vertically. Disadvantage of this arrangement, however, is the high complexity of this configuration and sensitivity in case of any failure of a subsystem, for example, the pivot mechanism of at least one rotor.
  • the rotors are designed quite large in their radial length extension, so that a safe landing in case of failure of a swivel mechanism is no longer successful.
  • this aircraft has to be reckoned with a high absolute and performance-specific energy consumption as well as a high level of noise during operation.
  • essential components of the aircraft may be damaged by leaking fragments of high energy for aviation safety, such as the landing gear arrangement and tail units.
  • the aircraft is also no longer able to fly.
  • Object of this invention is to find a safe arrangement possibility, embodiment and drive form for bladed rotors for an aircraft configuration, which is also in the sense of the open rotor concept for outside freely rotating drive rotors and for drive rotors high kinetic energy, and which is characterized that in the event of breakage of the rotor of the centrifugal force accelerated fragments more important for the safe operation of the flight component components of the aircraft, especially the pressure cabin, less likely to hit and damage.
  • Object of this invention is further to find solutions for propulsion rotors and propulsion systems on the aircraft, which allow a high bypass ratio up to today's usual high cruise speeds turbo-powered commercial aircraft, and to reduce the radiated noise during the drive and sustainable fuel efficiency of the powered aircraft to increase.
  • the invention is solved by an aircraft with:
  • a wing assembly W fixedly connected to the fuselage assembly F of the aircraft for generating a buoyancy carrying the aircraft, wherein the force component of the buoyancy generated perpendicularly to the fuselage longitudinal axis FA clearly outweighs the component along the fuselage longitudinal axis FA,
  • Dreiticiansanguin GA with several rotatably mounted wheels WE for statically determined support of the aircraft against the ground plane BO in ground-level operation, consisting of two structurally connected to the wing assembly W main landing gear GAM and structurally connected to the fuselage structure F Fugue leg GAA, wherein the nose gear leg GAA is arranged along the body longitudinal axis FA in the direction of flight FR in front of the main landing gear legs GAM, an engine arrangement E, comprising at least one engine, for generating an emergency service available to the aircraft,
  • At least one gear G for transmitting the drive power of at least one motor of the motor assembly E to at least one bladed rotor R
  • At least one bearing assembly S Dadurc h characterized in that at least one bladed rotor R with at least one position ran eleven S, the fuselage assembly F is rotatably mounted surrounding, and that at least this one bladed rotor R approximately in the direction of the fuselage longitudinal axis FA by at least one bearing assembly S axially fixed to the fuselage assembly F, and that at least this one bladed rotor R at the same m radial direction outwardly surrounds that portion of an at least partially cylindrically elongated fuselage assembly F of an aircraft annularly, which in one of the rotor RP views this rotor R, at least partially in the function of a pressurized cabin P can be printed and that at least in one of the rotor levels RP of the bladed rotor R no further for the un indirectly safe operation of the aircraft essential components of the aircraft are arranged, and that at least this one bladed rotor R, entlan Viewed in front of the wing assembly W and at the same time the geometrical axi
  • transmission F is kinematically coupled via transmission F, and can be set in rotation by the latter in order to produce a driving force driving the aircraft, whose force component in the direction of the longitudinal body axis FA clearly outweighs the force component perpendicular to the body longitudinal axis FA.
  • At least one bladed rotor R is arranged rotatably and rotatably supported at least partially printable fuselage assembly F, that the rotor R in one of its rotor planes RP, which in the region of its longitudinal extension along its rotor axis RA perpendicular to this are at least a portion of a pressurized cabin P, which lies within the fuselage assembly F surrounds.
  • the rotor viewed in one of its rotor planes RP, surrounds that cross-section of the fuselage assembly F which can be printed in this plane, at least in sections, according to a pressurized cabin P.
  • this section which can be imprinted on the function of a pressurized cabin P, is now really printed with an overpressure such that the pressurized cabin P within the fuselage arrangement has an overpressure to an area outside the pressurized cabin P inside or outside the aircraft, for example to that area outside the fuselage assembly F, in which at least one rotor R rotates in the fluid.
  • the bladed rotor R is rotatably mounted to the fuselage assembly F by at least one bearing assembly S and arranged fixed approximately in the direction of the fuselage longitudinal axis FA by at least one bearing assembly S.
  • Approximately to the direction of the fuselage longitudinal axis FA means in this context that the bladed rotor R can also have a conventional drive camber (approximately up to a maximum of 12 °) in relation to the fuselage longitudinal axis FA in a plane or spatially.
  • one or more rotors R are arranged along the body longitudinal axis FA on the fuselage assembly F, that in all of the rotor planes RP of a rotor, which is perpendicular to the rotational axis RP of the bladed rotor R and arranged within its longitudinal extent are arranged, no further necessary for the immediate safe flight of the aircraft important safety-critical components.
  • these too can not be damaged in a safety-critical manner.
  • Components that are important for the immediate safe execution of the flight may include, for example, the control surfaces, the spoilers, the high-lift system, the wing arrangement, fuel and flight control systems and important power distribution systems with their pipes and reservoirs, possibly also tail assemblies and landing gear arrangements and, of course, the pressurized cabin tanks and more engines count. In case of doubt, one can follow the licensing requirements of the aircraft.
  • An alternative method leads to a further embodiment of the invention.
  • a cone with the cone boundary starting from the extension limits of the bladed rotor R with a certain defined and suitable opening cone, e.g. of 2 x 15 °, which can also be referred to by appropriate statistics and investigations, define it as statistically probable that the trajectory of fragments leaving the rotor R under the effect of centrifugal forces at a suitable operating condition in case of failure , lies within exactly this cone.
  • the rotor R could be embodied radially outwardly in its circumference, at least over a certain angular range of the rotor R, in the circumferential direction encased by a cladding.
  • This panel can also be designed retractable, for example, so that it can be retracted or folded in the flow direction to the rear so that it fits aerodynamically as well as possible in a surrounding fuselage contour.
  • the fairing can advantageously increase the output thrust of the engine at low operating speeds, in particular at take-off.
  • this fairing can help to reduce the noise radiated downwards and sideways during take-off and landing by shielding. At cruising speed, however, falls. the power efficiency of the rotor higher when the shroud is retracted and sunk and no longer surrounds the bladed rotor R at least partially radially in the circumferential direction.
  • the sheathing could also be mounted directly on the outside of the rotor R and thus move in a rotating manner.
  • At least one bladed rotor R is designed such that it rotates unencumbered freely in the fluid. As a result, it achieves a very high efficiency in the propulsion conversion.
  • the bladed rotor R is designed overall so that its rotor blades B are adjustable in the setting angle to the flow. This makes it possible to use the high efficiency of the blasted rotor R over a wide speed range of the aircraft.
  • the setting angle change on the rotor blades is carried out by conventional mechanical devices in conjunction with actuators. Each rotor blade B can also be assigned a separate actuator.
  • two bladed rotors R are installed, which rotate in a mutually opposite direction of rotation. They form a rotor system together. In this way the rotor efficiency of the rotor system is improved by 7-8% in cruising speed and by up to 12% at takeoff relative to a single rotor. This contributes to a further advantageous improved efficiency of the aircraft.
  • the aircraft due to the rotation of the two rotors R, the aircraft does not experience a reaction moment on one side, since the reaction torques of the two rotors compensate each other.
  • one or more further pairs of rotors R rotating in opposite directions can be arranged at any desired fuselage position, the fuselage arrangement F, to these two bladed rotors R as well.
  • the aircraft does not unilaterally experience a reaction torque. This is compensated by the operation of two bladed rotors R with different directions of rotation relative to each other.
  • At least one bladed rotor R is arranged on the fuselage assembly F not only surrounding it, but in a certain way in relation to a 3-point landing gear assembly 6A with a nose gear GAA, which is provided with the statically determined Supporting the flight against the ground plane BO is used in ground-level operation is positioned so that at least one bladed rotor R is positioned on the fuselage assembly along the fuselage longitudinal axis FA between a nose gear GAA structurally connected to the fuselage assembly and a main landing gear assembly GAM, which to the airfoil assembly W is structurally tethered and includes two main landing gear legs.
  • the chassis legs of the main landing gear assembly GAM lie in the opposite direction of flight RFD, ie approximately in the flow direction, behind at least one rotor R so that the inflow of the rotor in the extended condition of the chassis assembly can not be disturbed by the main chassis legs.
  • the nose gear can be aerodynamically easier disguise so that this turbulence is minimized, for example, by a nose landing gear surrounding the surrounding lightweight fairing, or in addition by a wheel fairing similar to a so-called wheel - "slipper.”
  • This fairing can also be combined with a deflector known from other aircraft, eg, the main concorde landing gear, to prevent debris from being thrown up by the nose gear assembly GAA.
  • the traction arm is further away from the rotors R, and the aerodynamic trailing edge may also be outside the range of influence of the rotor surface at increased angle of attack during takeoff and landing.
  • a lowerable or extendable shielding protective device in the wake of a landing gear leg, especially the bow leg, in the direct line of sight between the chassis leg and bladed rotor R, a lowerable or extendable shielding protective device could be provided, which complicates that stirred up by the chassis foreign body below in the bladed rotor R can get damaged and this.
  • the extension can be done mechanically, electrically, hydraulically or with the help of the back pressure or weight-dependent.
  • a kind of fender could be attached to excellent wheels of the landing gear or on the landing gear flaps, which prevents foreign bodies can be whirled up by the chassis and get into the rotor.
  • At least one bladed rotor R is disposed between a nose landing gear assembly GAA and a main landing gear assembly GAM in the direction of the fuselage longitudinal axis FA on the fuselage assembly F such that in each of the rotor planes RP it views a bladed rotor R which largest outer diameter OD of this one bladed rotor R is smaller than that, given by the landing gear assembly GA, double minimum distance of the axis of rotation RA of this bladed rotor R from the ground plane BO. In this way, the bladed rotor R can safely rotate in all the usual phases of ground flush operation without coming into contact with the ground plane BO.
  • the bladed rotor R may also be advantageous to choose the largest outer diameter OD of a bladed rotor so that the bladed rotor R additionally has a safety distance to the ground plane BO.
  • This can make sense from a safety point of view, e.g. be chosen around the 60 cm, or also refer to the geometry of the wheels of the nose gear assembly GAA.
  • the distance of the bladed rotor R in one of the rotor planes RP of the bladed rotor R is influenced by the geometry of the chassis assembly GA, the position of the rotation axis RA of the bladed rotor R and the geometry of the fuselage assembly F in their cross section.
  • a very important further aspect of the invention is that at least one rotor R, preferably also two of them, is positioned in front of the latter in relation to the airfoil arrangement W in the direction of flight FR.
  • the aerodynamic flow of the rotor R can not be disturbed by the wing assembly W.
  • no turbulence of the lift-generating supporting surface W in the bladed rotor R which experience has shown that the radiated noise from the bladed rotor R would greatly increase.
  • in front of the support surface arranged bladed rotors R in their flow over their rotor surface a substantially homogeneous velocity distribution.
  • both the wake of the upper surface and the aerodynamic wake of the wing underside comprises, both of which have strong pressure andmatisdifferenzeh, what known edge vortex, which depart at the wing tips as wake turbulence proof.
  • the geometric axis of rotation RA of at least one bladed rotor R viewed in a plane immediately upstream of the wing assembly W and perpendicular to the axis of rotation RA of this bladed rotor R, and along the direction of the aircraft vertical axis VA of is located above the minimum vertical level of the upper surface of the aerofoil assembly W found in this plane, such that in a rotor plane RP of that one bladed rotor R the predominant aerofoil circumscribed by the rotor blades B of FIG Rotor R overstretched geometric rotor surface RAE, viewed from the ground plane BO, above the airfoil assembly W comes to rest so that the, on the opposite direction of flight RFD subordinate wing assembly W caused increase in pressure of the propeller jet at least this besch Aufelten rotor R predominantly above the wing assembly W takes place.
  • the propulsion generation according to the invention is not only low noise, but the propulsion necessary for propulsion simultaneously contributes to an increased lift generation on the wing assembly W at.
  • This further synergy effect according to the invention further advantageously increases the efficiency of the aircraft.
  • a rigid wing arrangement causes its buoyancy by causing a difference in velocity in the fluid between wing surfaces. Top and wing underside is induced, which causes a pressure difference between the top and bottom.
  • the lift generated on a conventional wing is due to a 2/3 by a negative pressure on the top and 1/3 by an overpressure on the underside of the wing. While this rule of thumb is not always true, as the exact geometry of the wing is a key factor, there is a strong tendency for the wing-top overspeed with respect to free flow to be greater than the velocity and thus pressure change on the wing underside.
  • the power introduced by the bladed rotor R and thus induced increase in velocity in the fluid, ie the increase in ram pressure, are introduced into the fluid predominantly above the support surface, which increases the lift.
  • a very significant advantage according to the invention also results from the fact that the rotor or rotors R in the direction of flight FR of the wing assembly W are connected upstream and also the guide Factory set LW of the wing assembly W is downstream in the flow direction.
  • the rotor R favored by the retractable landing gear of the landing gear assembly GA, for example, for maintenance, repair or replacement forwards, even in one piece, are pulled off over the fuselage. He does not have to be disassembled anymore.
  • the aircraft has a tail assembly, this consisting of a horizontal stabilizer assembly and a rudder assembly.
  • This tail assembly is in the direction of the fuselage longitudinal axis FA, seen in the direction of flight FR before and in the flow direction behind the Träg voman Aunt W installed so that both the horizontal stabilizer assembly and the rudder assembly at least absch italy arranged in direct wake of the propeller jet generated by at least one bladed rotor R. are.
  • the tail units are at least proportionately applied in the propeller jet of a rotor with an increased dynamic pressure and thus are more effective. Under certain circumstances, they can also be interpreted by their improved effectiveness with a smaller tail surface, so a total of low-resistance.
  • the direct wake of a propeller jet generated by a bladed rotor R can be determined by drawing straight lines downstream of the radial extension limits of a bladed rotor R parallel to the direction of flight FR.
  • these straight lines can also be oriented in the direction of the undisturbed aerodynamic inflow in front of the aircraft.
  • individual rotor blades B of at least one bladed rotor R can be adjusted independently of one another in their angle of adjustment for air flow, and also vary over the circumferential position of the rotor, so that a vectorization of the thrust results, with the control, for trimming , can be used to save on resistance and to reduce the noise of the aircraft.
  • the special reinforce inventive aspect of the additional lift increase by at least one bladed .Rotor R in this way additionally.
  • the setting angle of each rotor blade B is independent of the setting angles of other blades B, preferably by associated actuators, according to an Individual Blade Control concept, changeable.
  • the engine device E consists of several motors, each bladed rotor R is assigned to the power supply at least one particular engine of the engine device E.
  • the redundancy requirements can be met, and so, for example, a bladed rotor R of a motor of the motor assembly E and another bladed rotor R, which can also rotate in opposite directions, are driven by a second motor of the motor assembly E.
  • the engine assembly E consists of a plurality of engines, wherein from one of the rotor plane RP considered perpendicular to its axis of rotation RA of a bladed rotor R, these motors are arranged so that at least one engine in a region between the nose gear leg GAA and the right main landing gear of the main landing gear assembly GAM and at least a second motor in a region between the nose gear leg GAA and the left of the main landing gear legs of the main landing gear assembly GAM is arranged.
  • the motors of the motor assembly E seen in the spanwise direction of the wing assembly W, advantageously be installed close to the center of gravity, which due to the therefore smaller moments of inertia, the surfaces of the tail assembly LW can be made smaller to save resistance.
  • At least one motor of the motor assembly E in the opposite direction of flight RFD, that is seen approximately in the flow direction along the fuselage longitudinal axis FA arranged behind at least one bladed rotor R rotor system.
  • the motors of the motor assembly E can advantageously be installed close to the center of gravity in the direction of the fuselage longitudinal axis FA, wherein the surfaces of the tail assembly LW can be made smaller in order to save space due to the thus smaller moments of inertia of the aircraft.
  • At least one motor of the engine device E along the vertical axis VA of the aircraft, seen in the vertical direction from the ground plane BO, below the wing assembly W is arranged.
  • the motors of the motor assembly E in the direction of the vertical axis VA of the aircraft can advantageously be installed close to the center of gravity, due to the thus smaller moments of inertia of the aircraft, the surfaces of the tail assembly LW can be made smaller to save resistance.
  • a very low center of gravity also causes good handling properties of the aircraft in ground-level operation and especially during landing.
  • At least one bladed rotor R is driven by at least one gear G via at least one motor of the motor assembly E. It is ken n addressedd for the invention that this driving motor of the motor assembly E is housed outside the pressurized cabin P. This is done for reasons of practicability as well as for safety reasons, for example because of the installed volume, the cooling, the fire hazard and possibly the air supply and exhaust gas discharge.
  • the driving motor of the engine assembly E can also be arranged wholly or partially in the fuselage assembly F, but outside the pressurized cabin P.
  • the transmissions G for coupling at least one rotor R with at least one motor of the engine assembly E are also arranged outside the pressurized cabin P.
  • at least one bearing of the gear G is supported on the fuselage assembly F in order to structurally dissipate bearing forces there.
  • the gear G can also run at least proportionally within the fuselage assembly F.
  • the gear G mechanical, for example, non-positively, form-fitting, etc. work, also transmitted hydraulically, electrically or as a suitable hybrid form of these types of conversion their power.
  • a shaft turbine of the motor assembly E can drive a suitable generator that provides electrical power for an electric motor that at least proportionally drives at least a bladed rotor R in terms of power.
  • 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.
  • a bladed stator which converts the swirl energy of the air flow at least partially into further propulsion.
  • at least individual airfoils of the stator with respect to the free undisturbed air flow are designed to be adjustable, so that they also advantageous for vectoring the thrust, for control, trim, torque compensation after engine failure, to save on resistance or buoyancy increase of the Aircraft can be used with.
  • the blade adjustment as with a bladed rotor R, mechanically, electrically, hydraulically or by means of air loads via the back pressure.
  • At least one motor of the motor device E as a primary power output shaft gas turbine as a primary power output shaft gas turbine "be executed.
  • the shaft power can be transmitted, for example, via transmission to at least one bladed rotor R.
  • they could be kinematically connected to the bladed rotor so advantageously that they operate a large part of the flight mission in a particularly fuel-efficient manner in an advantageous stationary operating mode.
  • This advantage is further enhanced by the invention according to the invention for the blades B of the rotor R, a blade adjustment is provided for the thrust adjustment. This would allow the gas- bine largely stationary, because an adaptation and control of the thrust is possible not only on the speed adjustment of the gas turbine, but also on the blade adjustment.
  • At least one engine of the engine arrangement E is a gas turbine whose thermal efficiency is further increased in the sense of unconventional measures by recuperative means for compressor intermediate cooling, fuel mass preheating or exhaust gas heat exchange.
  • recuperative means for compressor intermediate cooling, fuel mass preheating or exhaust gas heat exchange can be used by the just described in the aircraft at least partially embedded layer, the space available there for such facilities that are highly space demanding, without that the resistance of the aircraft thereby significantly increases.
  • Previous problem according to the prior art it was that these facilities had to be arranged in the immediate vicinity of the engine and thus with in the engine nacelle. But this would have had to be significantly larger and disproportionately more resistant.
  • At least one motor of the motor device E is recessed in a region of the aircraft outside the pressurized cabin P, but at least partially recessed within a space which faces outward to the fluid through the outer skin of the airfoil assembly W and the is defined by the outer skin of the fuselage assembly F formed outer contour of the aircraft, and is arranged embedded in this space, that a component that dissipates 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 substantially not visible from the outside and that this component in its function of the motor mount outside the, by the outer skin of the wing assembly W and the outer skin of the fuselage assembly F formed outer contour formed substantially no own fluid-washed, resistive effective surface O is rdbar.
  • the engine or engine mounts no longer contribute to the resistance of the aircraft, and the engine nacelle can be omitted with its resistance-relevant flushed surface, except for air inlet and exhaust gas discharges.
  • the efficiency of the aircraft is advantageously higher. Due to the embedded position of the engines, the engine noise, especially the turbine noise, is effectively shielded from the outside.
  • one or more bladed rotors R can then be driven by means of remote shafts T, which are preferably parallel to the longitudinal axis of the aircraft, or by means of hydraulic power transmission systems.
  • the engines of the engine assembly E preferably as shaft engines or turboprocessing turbines for driving the bladed rotors
  • the engines of the engine assembly E are recessed and mounted with respect to all coordinate axes very close to the center of gravity. This in turn means for the aircraft, even with the rotors R, which are also mounted closer to the center of gravity that result in lower moments of inertia.
  • the structurally heavier T-tail can be omitted in favor of a conventional conventional cross-tail.
  • the exhaust gas stream may be led out in a further embodiment at the bottom of the hull stern so, moreover, that it in the region of the rear Rumpfeinschnü 'tion revives the boundary layer so energetic that an aerodynamic separation is reduced, which could reduce the resistance of the hull with.
  • the jet noise of the engines in the direction of the ground could also be shielded in the same train.
  • At least one engine of the engine assembly E has been housed in an unprinted area U of the aircraft, as shown by way of example, e.g. in a landing gear shaft. But it could also be usual unprinted areas U of the aircraft for it to be extended or new addition to be created.
  • motors of the motor assembly E are accommodated at least partially in the vicinity of the unprinted rear region of the fuselage assembly F. From here they could in turn drive the bladed rotors R arranged on the fuselage assembly F by means of long-wave shafts through a system for hydraulic power transmission. Any mixture of described arrangement options when using multiple engines and engines would be conceivable.
  • the bladed rotor R is placed around the fuselage assembly F of the aircraft. This results in the geometric rotor surface RAE from the circulation of the bladed rotor to the fuselage assembly F as a circular ring element of a relatively large inner diameter ID. This inner diameter ID is even larger than the Outer diameter of previous turbofan engines.
  • a circular ring element of a comparatively high inside diameter it is the case that with a specific blade span of the rotor blades B, a considerably larger geometric rotor surface RAE can be achieved than with a circular element with a blade span itself.
  • the largest possible rotor surface is in turn the prerequisite for a particularly fuel-efficient drive of the aircraft at low FPR and high mass flow.
  • the 'smallest inner diameter ID of at least one bladed rotor R is at least greater than the vertical height V of the pressurized cabin P and at the same time the vertical height V of this pressurized cabin P within the fuselage assembly F designed so constructively that there at least in sections, an upright gear is possible for the average passenger.
  • the geometry for the bladed rotor R is predetermined according to the invention such that a large geometric rotor surface RAE is formed with a small necessary span extension of the rotor blades B and thus sets a particularly efficient propulsion generation; and fuel efficiency on the aircraft.
  • the rotational speed of the rotor can preferably be optimally matched to this.
  • the transmission G in the drive train can be chosen in their transmission ratio so that not only for the bladed rotor R, but also for at least one engine of the motor assembly E optimal operating conditions with respect to the rotational speed and the moment are present, which increases the overall efficiency of the drive system.
  • the rotational speed with the increased thrust cross-sectional area or rotor area RAE could, for example, be chosen to be relatively low.
  • the thrust is then obtained mainly via a large mass flow, the blades B of the rotor R run at its hub as well as at their tip relatively slowly, which keeps the noise low.
  • the thrust generation of the bladed rotor R according to the invention is concentrated by the annular element on the radial section to 70-75% of the rotor diameter, which can generally produce the most effective and efficient thrust in rotors.
  • the by-pass ratio can be increased in one possibility according to invention thrust cross-sectional area of the sparged rotor R, which uriä the efficiency of the drive system further reduces the noise and increases the propulsion efficiency of the aircraft and thus its efficiency.
  • the applicable as a measure of the propulsion generation by the rotor blades B of a bladed rotor R in circulation swept geometric Rptor Structure RAE be made equal to or greater than the applicable as a measure of the hull resistance of the train and from bladed rotor R included transverse Cutting surface of the fuselage assembly F, which results in one of the rotor planes RP of this bladed rotor R perpendicular to its geometric axis of rotation RA.
  • this rotor must have in a plane perpendicular to the fuselage longitudinal axis FA at least half a fuselage diameter distance to the outer boundary of the fuselage assembly F of the aircraft so that it can rotate freely.
  • the fuselage assembly sidelined rotors now fails, there is a significant unfavorable moment by the lever arm in size at least one fuselage diameter.
  • the geometric rotor surface RAE swept in circulation by the rotor blades B of a bladed rotor R can be made equal to or larger than fifteen times the sum of the air inlet cross-sectional areas of all the ventilating motors of the motor arrangement E, which drive this bladed rotor R, the rotor blades B of this bladed rotor R are designed so aerodynamic that the aircraft flight speeds of at least Ma 0.76 at altitudes greater than 10000 m can be achieved. In this way it is possible with the drive arrangement to achieve higher bypass ratios in practice than is possible today with Geretefantriebmaschineen reach the bypass ratios of a maximum of 15: 1 or less. At the same time, the higher cruising speeds and cruising altitudes can be achieved for Schobofantriebwerke, which are overall larger than Turbopröpgetriebenen aircraft.
  • the arrangement advantages described also make it possible to safely use open-rotor and propeller rotors and drive systems on the aircraft. This also allows higher bypass ratios to be realized at the comparatively high cruise speeds of today's turbofan engines. This reduces the fuel consumption compared to today operated aircraft, sustainable.
  • the arrangement according to the invention of the rotor, annularly surrounding the hull gives the possibility of greater freedom in the aerodynamic design of the bladed drive train.
  • rotors and retroactively also to use the drive system the overall favor a more fuel-efficient and low-noise design of the drive system and allow a higher propulsion efficiency for the aircraft.
  • the effective thrust circle surface could be selected to be larger overall in terms of area. For a certain airspeed, it then results, in order to achieve a given thrust, to be able to reduce the speed distribution given to the fluid by the bladed rotor R, and thus to be able to reduce the pressure ratio of the bladed rotor R.
  • the load of the individual blades of the rotor can be reduced, for example, by the pitch or angle of the blades is reduced, or produce them by a modified profiling aerodynamically less lift per length.
  • there is a lower thrust area load, or in this case a lower shear ring area load less rotor noise and increased fuel efficiency.
  • the remaining parameters such as the pressure ratio of the bladed rotor R, the angular velocity of the rotation and the slope of the blading, with respect to a reference engine, left approximately the same with a larger thrust cross-sectional area results in a higher thrust for the aircraft.
  • the blades of the rotor R can be preferably even in feathered position, for example, in Nullauftriebsraum brought, where the additional resistance can be significantly reduced by this feathering.
  • the adjustment can be organized automatically in case of engine failure. It is another important performance advantage over conventional engines.
  • the blades can in a further embodiment, also by folding back, for example, to the fuselage assembly F in feathered position be brought, similar to a folding propeller. In the case of multi-engine airplanes, this would be an advantage in terms of performance in terms of performance.
  • the distribution of the rotor blades B e.g. in their setting angle, also be designed so that a thrust reversing action of the bladed rotor R for braking on the ground and possibly also in the air is possible.
  • a thrust reversing action of the bladed rotor R for braking on the ground and possibly also in the air is possible.
  • the bladed rotor R also often referred to as a "propulsor" in its mode of operation, is such that, as previously stated, it can rotate by means of a suitable bearing arrangement S on the aircraft with respect thereto and by a suitable design of the blades B
  • the bladed rotor thus provides thrust generation with the effect of a propulsion force on the aircraft
  • This propulsion force driving the aircraft is oriented in this way, the force component thereof in the direction of the fuselage longitudinal axis FA opposite
  • the bladed rotor R takes over the production of propulsion in substantially horizontal flight direction and thus essentially does not contribute to its function, as in a rotorcraft, to generate the relevant buoyancy for the aircraft.
  • the bladed rotor in one possible embodiment can be designed such that the blades B of the rotor R are characteristic of aerodynamic parameters such as thrust cross-sectional area, circle area loading, rotational angular velocity, blade span, blade depth, relative blade thickness , Twisting, tapering, Schaufelfeilung, profiling, the number of blades and the overall appearance are designed similar to the blades of a fan.
  • the bladed rotor R can be embodied in a further embodiment so that the blades B of the rotor with respect to characteristic aerodynamic parameters, such as the thrust cross-sectional area, circle (ring) surface load, rotational angular velocity, blade span.
  • characteristic aerodynamic parameters such as the thrust cross-sectional area, circle (ring) surface load, rotational angular velocity, blade span.
  • Shovel depth, relative thickness of the blade, twisting, tapering, Schaufelpfeilung, profiling, the number of blades and the overall appearance are designed similar to the blades of a Propfans.
  • the bladed rotor R may be designed such that the blades (B) of the rotor have characteristic aerodynamic parameters, for example thrust cross-sectional area, circle (ring) area loading, rotational angular velocity, blade span, blade depth, relative thickness Shovel, twisting, tapering, Schaufelpfeilung, profiling, the number of blades and the overall appearance are designed similar to the blades of an air screw.
  • characteristic aerodynamic parameters for example thrust cross-sectional area, circle (ring) area loading, rotational angular velocity, blade span, blade depth, relative thickness Shovel, twisting, tapering, Schaufelpfeilung, profiling, the number of blades and the overall appearance are designed similar to the blades of an air screw.
  • characteristic aerodynamic parameters for example thrust cross-sectional area, circle (ring) area loading, rotational angular velocity, blade span, blade depth, relative thickness Shovel, twisting, tapering, Schaufelpfeilung, profiling, the number of blades and the overall appearance are designed similar to the blades of an
  • the airfoils B in outward radial extension are designed so that they can act less propulsive generating in a span position near its hub as at a span position, which is approximately 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. 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 by means 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.
  • rotor blades B could be designed differently from each other, for example, have a different span to favorably influence the noise. In rotary operation, this still results in an annular thrust cross-sectional area of the bladed rotor.
  • the bladed rotor R is also according to the invention suitable for being kinematically attachable to an engine assembly of the aircraft by suitable features as part of a transmission such that the rotor R can generate propulsion driving the aircraft by its rotation.
  • it can be attached to the engine arrangement of the aircraft as part of a mechanical transmission.
  • it has as a feature, for example, a suitable toothing along one of its sides or along its circumference.
  • This toothing can preferably be arranged on the inside or on the outside, the sides or even on any inclined plane and also use a toothed form, which ensures a reliably high power transmission, for example, a helical toothing or a toothing on which bevel gears can intervene.
  • the rotor can then be coupled, for example, for driving on at least one toothed wheel driven by at least one motor of the motor arrangement E.
  • the bladed rotor could also be embodied as part of a hydraulic transmission in a further embodiment. In this case, this transmission should have a particularly advantageous high efficiency. If it coincides at least proportionally with an annular electric motor, it thus forms the part of an electrical, sometimes even an electronic transmission.
  • counter rotating bladed R and drive systems can be used by different power control also to control the aircraft with about the longitudinal axis when rolling.
  • the at least partial roll control or trimming could take place via different rotational speeds, torque specifications or by varying the position of the blading of the rotors and possibly the stators.
  • the bladed rotors R can be driven by one or more motors of the fylotorenan Aunt E via remote shafts T or systems for hydraulic power transmission, which are housed in the unprinted areas U at least proportionately within the aircraft.
  • the bearing arrangement S can be carried out preferably by an arrangement of rolling bearings, plain bearings or a suitable combination of both types of bearings.
  • the bearing arrangement S of the rotors could also be formed by hydraulic bearings, for example by hydrostatic bearings.
  • magnetic bearings could be used in the storage of the rotor. Both types of bearings mentioned above offer the possibility of being able to additionally dampen the noise and vibrations to the aircraft through active control.
  • piezoelectrically variable and electrically controllable damper which .An active vibration and vibration damping allow.
  • Magnetic bearings also ensure a very low bearing friction, which has an advantageous effect on the efficiency of the drive system.
  • Magnetic bearing e.g. Superconducting magnetic bearings or electrodynamic magnetic bearings need, according to the current state of the art, a certain 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 high support width in the storage.
  • the drive unit formed by the bladed rotor R and the associated bearing and guide unit, may additionally comprise panels O, e.g. also to the fuselage assembly F of the aircraft to aerodynamically additionally influence the flow, or to increase the aerodynamic quality of the drive system,;
  • the storage in a further embodiment can also take place together with an electric motor, for example so that the bladed rotor R advantageously simultaneously forms the rotor of the electric motor.
  • the storage of the common rotor can in particular be such that the storage and the electric motor further component components share together, for example, magnets, coils or a cooling system.
  • a rotor R could also be part of a "stator" of at least one electric motor and a further rotor R part of a rotor of an electric motor form such that the rotors move in opposite directions of rotation with respect to each other and transmit no torque to the fuselage assembly F becomes.
  • the now aerodynamically clean wing of the wing assembly W may preferably be equipped with additional aerodynamic drag saving concepts, e.g. with laminar retention 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 more technically effective.
  • An inventive advantage is that the bladed rotor R rotatably mounted at low additional weight preferably structurally can be connected directly to the fuselage assembly F.
  • a plurality of motors of the motor assembly E may drive a bladed rotor R in parallel.
  • the rotor in fail-safe design. tion on the aircraft and thus meet the redundancy requirements of the drive system so that several engines are provided, which could drive only one rotor in parallel R.
  • the architecture can also be chosen such that, in the event that one of the motors of the motor assembly E fails, the remaining motor or motors continue to provide enough drive power to drive the bladed rotor R so in critical flight phases of the aircraft so that he Effectively meets the performance requirements of the aircraft.
  • At least one of the motors of the motor assembly E could be selectively coupled in or out by a clutch for driving the rotor R.
  • a clutch for driving the rotor R For example, in the event of a malfunction of a motor, this could be decoupled from the drive system. If the power requirement increases, preferably a further motor could be coupled in.
  • the rotors can be coupled on the ground as needed, sometimes for rolling.
  • At least one of the engines of the engine assembly E can be used not only as an internal combustion engine, e.g. be designed as a gas turbine or piston engine, but also as an electric motor.
  • motors of the motor assembly E which are housed for example in the area of the wing-fuselage arrangement or in the chassis shaft, are very well 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 the bladed rotors R and possibly for the longwave T, for example, below an unprinted aerodynamic fairing on the outside Hull outside, lying on this lying, may be appropriate.
  • At least one electric motor of the motor assembly E also temporarily at least temporarily provide additional drive power, preferably in flight phases that require at least short-term increased demand thrust of the aircraft, for example, during takeoff, the climb or in the event of an engine failure.
  • annular electric motor as part of the engine assembly E, substantially concentric with the bladed rotor R surrounding the hull F, could also be arranged to drive the bladed rotor R.
  • This electric motor could also be partially embedded in the fuselage assembly or also form a Bestanteil the fuselage assembly F. From the wind energy generation powerful annular electric motors are known, which are used there as a generator and which would basically be operated as motors. Current research assumes that the mass-specific performance weight of these components can be improved by a significant factor in the future.
  • This arrangement also offers the advantage that an electric motor, which is in kinematic interaction with the bladed rotor, is effectively cooled during operation by the circulating air in the cruise about -60 ° C external environment.
  • At least one air-breathing engine of the engine assembly which drives at least one bladed rotor R produces, in addition to the propulsive force of at least one rotor R, further propulsive force in the form of thrust by the emission of exhaust gases, the force component also being at this propulsive force Longitudinal direction of the axis of rotation RA against the force component perpendicular to the axis of rotation RA clearly outweighs.
  • the bladed rotor R becomes inventive in its surface-specific thrust load relieved, so that he can work more effectively in certain operating conditions of the aircraft, for example, at takeoff. It thus follows, similar to a turbofan engine, on aircraft level an engine assembly, wherein the sheath stream envelops the hull assembly and the shaft turbine with the core flow also generates thrust, but on a larger dimension level and level of the possible bypass ratio.
  • the largest outer diameter OD of at least one bladed rotor is made equal to or smaller than twice the maximum longitudinal extent LGAM of a chassis leg together with wheels WE of the main undercarriage assembly GAM.
  • At least one bladed rotor R is at least partially surrounding the fuselage F, which advantageously degrades the additionally induced by the bladed rotors speed of the fluid and at least partially converts into additional pressure, in addition, by the curvature due to the relative positive pressure in the flow, which creates a thrust-generating effect on the aircraft (similar to the principle of thrust recovery).
  • This diffuser may also result from a change in the fuselage diameter of the fuselage assembly F or through its fairings.
  • an electric motor is used in the drive system which is kinematically connected to the bladed rotor R, it could be preferably, e.g. be operated in suitable operating phases of the aircraft, in a further embodiment also as a generator.
  • the invention thus lays the foundation for realistic architectures of future hybrid-electric propulsion systems in aircraft.
  • the bladed rotor R can be used in a further embodiment with the acceleration of the fuselage boundary layer, this leads to energy, which can thus further reduce the resistance of the fuselage and the aircraft.
  • the fuselage boundary layer could be at least partially suctioned off upstream of at least one rotor R, at least partially, in order to avoid unfavorable interaction with the rotor R.
  • the nose landing gear arrangement GAA can be extended and retracted independently of the main landing gear arrangement GAM.
  • bladed rotors R can be deducted according to the invention forward over the fuselage assembly F for replacement, maintenance or repair without having to be disassembled in its scope.
  • the thrust cross-sectional area or the cross-sectional area of the besehaufelten rotor in terms of a rotor surface or a Rotornikring Structure that surface of the bladed rotor is understood that arises in the rotary operation by the sweeping of the airfoils B in a plane that normal to the axis of rotation RA of the rotor R.
  • turbo-prop turbine in this document, by function, is the shaftless gas turbine engine unit without propellers. This may already have a reduction gear in a possible Aiis Equipmentsform internally.
  • RFD designates a direction which is directed opposite to the direction of flight of the aircraft, that corresponds approximately to the flow direction of the fluid.
  • the patent application seeks protection for a non-vertically takeoff and landing surface aircraft, preferably the genre of a commercial aircraft.
  • E engine assembly formed by at least one engine
  • Hull arrangement pressurized cabin
  • other A. shape pressurized cabin printed with overprint
  • RP rotor plane that 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 extension greater than the spatial extent of the train, to determine the impact area of the bladed rotor R
  • V Vertical cabin height within the fuselage layout
  • Figure 1 is a highly simplified exemplary representation of the side view of a future commercial aircraft with two rotatably mounted bladed rotors R, each rotor R can be rotationally driven by an associated shaft turbine of the motor assembly E via an associated remote shaft T and, wherein the shaft turbines within an unprinted Area U in the vicinity of the wing-fuselage arrangement, there are partially sunk in the plane lying inside, arranged, and wherein in this representation, the second Welienturbine with its remote shaft is not directly visible because it, due to the representation, behind the first Drive system is arranged concealed in the spanwise direction;
  • FIG. 2 is a highly 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 can be rotationally driven by a respective shaft shaft of the motor assembly E via an associated remote shaft T via gear and, wherein the shaft turbines located within an unprinted area U near the wing-fuselage assembly, there partially sunk inside the aircraft;
  • Figure 3 is an overview of previously customary printed areas P and unprinted areas U within the fuselage assembly F of a transport aircraft;
  • FIG. 4 shows the exemplary secure arrangement of two bladed rotors R along the fuselage assembly F on the aircraft according to an alternative detection method such that within the excursion cones, symbolized by virtual torus bodies, do not have any desired opening angles emanating from the bladed rotors for the safe flight necessary other components of the aircraft are;
  • Figure 5 shows a greatly simplified exemplary rotatable arrangement of a rotor R according to the invention with a bearing assembly S, the profile section of the fuselage assembly surrounding annularly F, wherein within the cross section of the fuselage assembly F according to a pressure cabin P printable portion is provided by its vertical height V so high it is stated that an average passenger can stand upright in it;
  • FIG. 6 shows a profile section of an exemplary embodiment of a bearing and guide unit with a bearing arrangement S for rotatably supporting the drive rotor R with simultaneous drive-effective removal of the axial forces and kinematic coupling of the bladed rotor R via transmission to a spring shaft T,
  • FIG. 6 shows a profile section of an exemplary embodiment of a bearing and guide unit with a bearing arrangement S for rotatably supporting the drive rotor R with simultaneous drive-effective removal of the axial forces and kinematic coupling of the bladed rotor R via transmission to a spring shaft T
  • FIG. 7 shows a greatly simplified exemplary kinematic coupling of a bladed rotor R according to the invention via gearboxes to a plurality of motors of the motor assembly E, wherein the bladed rotor R has an internal toothing, to which a further toothed wheel, which is connected kinematically to a motor of the motor assembly E, here by way of example as a hybrid-electric embodiment of the drive system, wherein a motor of the engine assembly E is designed via a clutch and disengaged disengageable; wherein an engine is designed as a gas turbine and the further motor as an electric motor;
  • Figure 8 shows a highly simplified exemplary sketch of the possible arrangement of a bladed rotor, shown in a plane at a fuselage length position at the height of the airfoil assembly 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 of at least one bladed rotor equal or smaller than twice the maximum longitudinal extent of a chassis leg together with wheels WE of the main landing gear arrangement GAM;
  • FIG. 9 shows a greatly simplified exemplary sketch of a rotor R bladed with rotor blades B, left side view and right front view, and illustrates the definition of rotor planes RP in the region of the longitudinal extent of the bladed rotor R along its rotational axis RA Rotor planes RP are normal to the rotation axis RA;
  • FIG. 1 shows in a greatly simplified form an exemplary embodiment of a future commercial aircraft according to the invention in a side view.
  • here are two bladed rotors R so arranged at least partially printable fuselage assembly rotatably mounted surrounding this, that each rotor R in one of its rotor planes RP, which in the region of its longitudinal extent along its rotor axis RA perpendicular to this are arranged, at least part of a pressurized cabin P, which lies within the fuselage assembly F, surrounds.
  • each rotor R seen in one of its rotor planes RP, here surrounds that cross-section of the fuselage assembly F which can be printed in this plane at least in sections according to a pressurized cabin P.
  • the bladed rotors R are rotatably mounted to the fuselage assembly F by at least one bearing assembly S (as exemplified in Figure 6) and arranged approximately in the direction of the fuselage longitudinal axis FA by at least one bearing assembly S (as exemplified in Figure 6).
  • the axes of rotation RA coincidentally coincident here in these two rotors R, in the sense of a conventional drive camber at a small angle with respect to the body longitudinal axis FA inclined.
  • the two rotors R are inventively arranged along the fuselage longitudinal axis FA on the fuselage assembly F, that in all of the rotor planes RP of each of the two rotors, which are perpendicular to the axis of rotation RA of the bladed rotor R and within its longitudinal extent, no further for the immediate safe flight of the aircraft necessary important safety-critical components are arranged. Thus, in the event of a broken rotor R, these too can not be damaged in a safety-critical manner.
  • both bladed rotors R are designed so that they rotate freely in the fluid uncoated. As a result, they achieve very high efficiency in the propulsion conversion.
  • the two bladed rotors R are advantageously installed in this application example so that they rotate with an opposite direction of rotation. They form a rotor system together. In this way, the rotor efficiency of the rotor system is improved.
  • the bladed rotors R are arranged on the fuselage assembly F these not only surrounding, but in a certain way compared to a 3-point suspension arrangement GA, a total of several rotatably mounted wheels WE, and with e, inem nose gear GAA, where GA with the statically determined Supporting the flight against the ground plane BO in ground-level operation, positioned so that both bladed rotors R are positioned on the fuselage assembly along the fuselage longitudinal axis FA between a nose gear GAA structurally attached to the fuselage assembly and a main landing gear assembly GA, the latter to the airfoil assembly W is structurally tethered and includes two main landing gear legs.
  • the applicable as a measure of the propulsion generation of the rotor blades B of a bladed rotor R in circulation swept geometric rotor surface RAE equal or larger than the applicable as a measure of the trunk resistance of the aircraft and the bladed rotor R enclosing cross-sectional area of the fuselage assembly F. which results in one of the rotor planes RP of this bladed rotor R perpendicular to its geometric axis of rotation RA.
  • both rotors R are positioned in front of the latter in relation to the wing arrangement W in the direction of flight FR.
  • the aerodynamic flow of the rotor R can not be disturbed by the wing assembly W.
  • no turbulence of the lift generating wing W in one of the bladed rotors R which would experiencely increase the radiated from the bladed rotors R noise greatly.
  • the bladed rotors R arranged in front of the wing undergo a largely homogeneous velocity distribution in their inflow over their rotor surface.
  • the rotors R are arranged upstream in the direction of flight FR of the wing assembly W and also the tail assembly LW of the wing assembly W is arranged downstream in the flow direction.
  • the rotor R favored by the retractable landing gear of the landing gear assembly GA, for example, for maintenance, repair or replacement to the front, even in one piece, be pulled over the fuselage. He does not have to be disassembled anymore.
  • the aircraft has a tail assembly, this consisting of a horizontal stabilizer assembly and a rudder assembly.
  • this tail assembly is installed according to the invention upstream and downstream of the wing assembly W such that both the tailplane assembly and the rudder assembly are at least partially in the direct wake of the propeller jet generated by at least one bladed rotor R. are arranged.
  • the tail units are thereby, at least proportionally, lying in the propeller jet of a rotor, subjected to an increased dynamic pressure and thus are more effective.
  • the engine device E consists of two motors, each bladed rotor R being associated with the power supply E of at least one of these two dedicated motors.
  • an engine of the engine assembly E drives a first bladed rotor R via a remote shaft T and via a gear G.
  • a second motor of the engine assembly E also drives the second bladed rotor R via an associated remote shaft T and via gears.
  • the longwave T are on opposite sides of the fuselage assembly F, partially stored at this, arranged.
  • the gear G is partially integrated on the rotating rotor ring.
  • This rotor ring of the bladed rotor has here an internal toothing, by another gear engages, which is in kinematic operative connection with the remote shaft T, which is driven by a motor of the motor assembly E.
  • the kinematic operative connection can be done directly, for example, via a shaft-hub connection or generally via other gears of the transmission G.
  • the rotor ring thus simultaneously forms the larger gear of a reduction gear between bladed rotor R and motor assembly E.
  • the two motors of the motor assembly E are bladed in one of the rotor plane RP perpendicular to its axis of rotation RA of a.
  • Rotor R is arranged so arranged that at least one motor in a region between the nose gear leg.GAA and the right main landing gear GAM and at least a second motor in a region between the nose bridge GAA and the left of the main landing gear legs of the main landing gear assembly GAM is arranged.
  • the motors of the motor winding E can advantageously be installed close to the center of gravity.
  • the two motors of the motor assembly E are arranged in this embodiment, in the opposite direction of flight RFD, so seen approximately in the flow direction, along the fuselage longitudinal axis FA behind at least one bladed rotor R rotor system.
  • the motors of the motor assembly E in the direction of the fuselage longitudinal axis FA can advantageously be installed close to the center of gravity.
  • the motors of the engine device E along the vertical axis VA of the aircraft seen in the vertical direction from the ground plane BO, below the wing assembly W are arranged.
  • the motors of the motor assembly E in the direction of the vertical axis VA of the Airplane advantageous be installed close to the center of gravity.
  • the two motors of the engine E are designed in this example as a primary shaft output donating gas turbine.
  • Each of the two bladed rotors R is here driven via at least one gear G via an associated engine of the engine assembly E, wherein these driving motors of the engine assembly E proportionately in the fuselage F, but outside the pressure cabin P, are arranged.
  • the air inlets are suitably led far out of the unprinted area U and also the engine exhaust gases are discharged out of this space via guides.
  • the tail assembly is designed as a conventional cross-tail, he external noise of the propeller turbine is greatly reduced by the sunk arrangement in the aircraft.
  • the turbos power turbines that drive the bladed rotors, additional thrust in the sense of propulsive force for the aircraft.
  • the exhaust gas jet (here indicated by a gray dashed arrow) can be brought out at the underside of the fuselage tail in such a way that it can energize the boundary layer in the region of the rear fuselage neck in such a way that an aerodynamic detachment is reduced Resistance of the hull with could minimize.
  • both bladed rotors R are arranged between a nose landing gear assembly GAA and a main landing gear assembly GAM toward the fuselage axis FA on the fuselage assembly F such that each of these rotor blades RP views each of these bladed rotor Rs bladed rotor R is ner, as, with specified by the landing gear assembly GA, double minimum distance of the axis of rotation RA of this bladed rotor R of the ground plane BO.
  • the bladed rotor R can safely rotate in all the usual phases of ground flush operation without coming into contact with the ground plane BO.
  • the largest outer diameter OD of the bladed rotors is selected so that the bladed rotor R additionally has a safety distance of about 60 cm to the ground plane BO.
  • FIG. 2 shows a highly simplified form of an exemplary embodiment of a future commercial aircraft according to the invention in the front view.
  • here are two bladed rotors R so arranged an at least partially druckbäre fuselage assembly F rotatably mounted surrounding this, that each rotor R in one of its rotor planes RP, which arranged in the region of its longitudinal extent along its rotor axis RA perpendicular to this are at least a portion of a pressurized cabin P, which lies within the fuselage assembly F surrounds.
  • each rotor R seen in one of its rotor planes RP, here surrounds that cross-section of the shirring arrangement F which can be printed in this plane at least in sections according to a pressure booth P.
  • the bladed rotors R are rotatably mounted to the fuselage F by at least one bearing assembly S (as shown by way of example in Figure 6) and arranged approximately in the direction of the fuselage longitudinal axis FA by at least one bearing assembly S (as exemplified in Figure 6).
  • the axes of rotation RA which coincide coincidentally here in these two rotors R, in the sense of a conventional drive camber inclined at a small angle with respect to the body longitudinal axis FA executed.
  • the two rotors R are inventively arranged along the fuselage longitudinal axis FA on the fuselage assembly F, that in all of the rotor planes RP of each of the two rotors, which are perpendicular to the axis of rotation RA of the bladed rotor R and within its longitudinal extent, no further for the immediate safe flight of the aircraft necessary important safety-critical components are arranged. Thus, in the event of a broken rotor R, these too can not be damaged in a safety-critical manner.
  • both bladed rotors R are here designed such that they rotate ungemantelt f (rei in the fluid according to a special characteristic of the invention. As a result, reach these very high efficiency in the propulsion conversion.
  • the two-bladed rotors R are advantageous in this application example, as installed, In this way the rotor efficiency of the rotor system is improved
  • the bladed rotors R are arranged on the fuselage assembly F, not only surrounding them, but in a certain way opposite a rotor assembly 3.
  • Point landing gear assembly GA in total with several rotatably mounted wheels WE and with a nose gear GAA, where GA with the statically determined Supporting the flight against the ground plane BO is used in ground-level operation, positioned so that both bladed rotors R are positioned on the fuselage assembly along the fuselage longitudinal axis FA between a nose gear GAA structurally attached to the fuselage assembly and a main landing gear assembly GAM, the latter to the airfoil assembly W is structurally tethered and includes two main landing gear legs.
  • the applicable as a measure of the propulsion generation of the rotor blades B of a bladed rotor R in circulation swept geometric rotor surface RAE equal or larger than the applicable as a measure of the hull resistance of the train and enclosed by the bladed rotor R cross-sectional area of the fuselage assembly F. which results in one of the rotor planes RP of this bladed rotor R perpendicular to its geometric axis of rotation RA.
  • both rotors R positioned with respect to the wing assembly W in the direction of flight FR in front of this.
  • the aerodynamic flow of the rotor R can not be disturbed by the wing assembly W.
  • no turbulence of the lift generating wing W in one of the bladed rotor R which would experience the radiated radiated from the bladed rotors R noise would greatly increase.
  • the rotors R are upstream in the direction of flight FR of the wing assembly W and also the tail assembly LW of the wing assembly W is arranged downstream in the flow direction.
  • the rotor R favored by the retractable landing gear of the landing gear assembly GA, for example, for maintenance, repair or replacement forwards, even in one piece, are pulled off over the fuselage. He does not have to be disassembled anymore.
  • the aircraft has a tail assembly, this consisting of acitenleittechniksari onion and a rudder assembly.
  • This tail assembly is in the direction of the fuselage longitudinal axis FA, seen in the direction of flight FR before and in the flow direction behind the wing assembly W so installed according to the invention that both the horizontal stabilizer assembly and the rudder assembly, at least partially, in the immediate wake of at least 'a bladed rotor R generated Propeller beam are arranged.
  • the tail units are thereby, at least proportionally lying in the propeller train of a rotor, subjected to an increased dynamic pressure and thus are more effective.
  • the engine device E consists of two motors, each bladed rotor R being associated with the power supply E of at least one of these two dedicated motors.
  • a second motor of the motor assembly E also drives via an associated remote shaft T and. Gear on the second bladed rotor R.
  • the longwave T are on opposite sides of the fuselage assembly F, partially stored at this, arranged.
  • the gear G is partly integrated on the rotating rotor ring.
  • This rotor ring of the bladed rotor has here an internal toothing, by another gear engages, which is in kinematic operative connection with the remote shaft T, which is driven by a motor of the motor assembly E.
  • the kinematic operative connection can be directly, for example via a shaft-hub connection or generei! also via other gears of the transmission G, done.
  • the rotor ring thus forms at the same time the larger gear of a reduction gear between bladed rotor R and motor assembly E.
  • the two motors of the motor assembly E are freighted in one of the rotor plane RP perpendicular to its axis of rotation RA of a bladed rotor R, arranged so that at least one motor in one Area between the nose gear leg GAA and the right main landing gear of the main landing gear assembly GAM and at least a second motor in a region between the nose gear leg GAA and the left of the main chassis Beiiie the main landing gear assembly GAM is arranged.
  • the motors of the motor assembly E can advantageously also be installed close to the center of gravity.
  • the two engines of the Engine assembly E in this embodiment in the opposite direction of flight RFD, so seen in approximately the flow direction along the fuselage longitudinal axis FA arranged behind at least one rotor blade R rotor system formed.
  • the motors of the motor assembly E in the direction of the fuselage longitudinal axis FA can advantageously be installed close to the center of gravity.
  • the motors of the engine device E along the vertical axis VA of the aircraft, seen in the vertical direction from the ground plane BO, below the wing assembly W are arranged.
  • the motors of the motor assembly E in the direction of the vertical axis VA of the aircraft can be advantageously installed close to the center of gravity.
  • the two motors of the engine device E are designed in this example as a primary shaft output donating gas turbine.
  • Each of the two bladed rotors R is here driven via at least one gear G via an associated engine of the engine assembly E, wherein these driving motors of the engine assembly E proportionately in the fuselage assembly F, but outside the pressurized cabin P, are arranged.
  • the air inlets are suitably led far out of the unprinted area U and also the engine exhaust gases are discharged out of this space through guides.
  • the turboshaft turbines which drive the bladed rotors generate additional thrust in the sense of a propulsion force for the aircraft.
  • the exhaust gas jet (indicated here by a gray dashed arrow) can also be found in of this exemplary embodiment are brought out at the bottom of the fuselage tail so that it can energize the boundary layer in the region of the rear fuselage so energetically that an aerodynamic detachment is reduced, which could reduce the resistance of the fuselage.
  • both bladed rotors R are disposed between a nose gear assembly GAA and a main landing gear assembly GAM in the direction of the fuselage longitudinal axis FA on the fuselage assembly F so that in each of the rotor planes RP, viewed each of these bladed rotors R, the largest outer diameter OD of the bladed rotor R is smaller than that, given by the landing gear assembly GA, double minimum distance of the axis of rotation RA of this bladed rotor R of the ground plane BO.
  • the bladed rotor R can safely rotate in all the usual phases of ground flush operation without coming into contact with the ground plane BO.
  • the largest outer diameter OD of the bladed rotors is selected so that the bladed rotor R additionally has a safety distance of about 60 cm from the ground plane BO.
  • Figure 3 is an overview of previously customary printed areas P and unprinted areas U within the fuselage assembly F of a transport aircraft; wherein the lying in the fuselage assembly F parts, which are shown in white, can be printed as a printing booth P, according to the invention, also other unprinted areas created or existing existing can be increased.
  • Figure 5 is a greatly simplified, exemplary rotatable arrangement of a rotor R according to the invention with a bearing assembly S, the profile section of the fuselage assembly surrounding annularly, wherein within the cross section of the fuselage assembly F according to a pressure cabin P printable portion is provided by its vertical height V so high that an average passenger can stand upright; wherein only the upper part of the fuselage cross section is printed here corresponding to a pressure cabin, while the lower is delimited from the upper part by a kind of floor assembly.
  • the bladed rotor has a smallest inner diameter ID and a largest outer diameter OD.
  • Figure 6 shows a profile section of an exemplary and highly simplified design of a bearing and guide unit with a bearing assembly S for rotational storage of the ⁇ ntriebsrotors R while driving efficient dissipation of axial forces and kinematic coupling of the bladed rotor R via gearbox to a remote shaft T, the shaft power on the remote shaft T engages via bevel gears in a toothed ring according to a gear G, wherein the internally toothed ring is part of the bladed rotor R surrounding the body assembly F;
  • FIG. 7 shows a greatly simplified exemplary kinematic coupling of a bladed rotor R according to the invention via gearboxes to a plurality of motors of the engine arrangement E, wherein the inspected feite rotor R has an internal toothing on which a further gear, which is connected to a motor of the motor arrangement E, kinematic acts, here for example as a hybrid e ektharide of the drive system, wherein a motor of the motor assembly E via a coupling a and is executed disengageable; wherein a motor is designed as a gas turbine and the further motor as an electric motor, wherein representatively simplified only some of the rotor blades B are shown on the rotor;
  • FIG. 8 shows a greatly simplified exemplary sketch of the possible arrangement of a bladed rotor, shown in a plane at a fuselage length position at the level of the airfoil assembly W downstream of the main landing gear assembly GAM perpendicular to the axis of rotation of a bladed rotor, seen here in the direction of flight FR, the largest Outer diameter OD of at least one bladed rotor is equal to or smaller than twice the maximum longitudinal extent LGAM of a chassis leg together with wheels WE of the main landing gear arrangement GAM.
  • an arrangement of the main landing gear GAM is possible in such a way that the two landing gear legs attached to the wings disturb neither the flow nor the outflow of a rotor R.
  • the Fahftechnik as indicated here, pivoted inward and retracted.
  • the available space then extends through this geometric arrangement also from the length for the retraction of both legs.
  • the rotor can not geometrically touch the ground plane.
  • it is also represented by way of example transferable in this plane that in this plane the axis of rotation RA of a rotor along the aircraft vertical axis VA, seen from the ground plane BO, above a vertical minimum MV of the upper side of the wing assembly W comes to rest, so, that the predominant part (here indicated by dashed lines) of the area RAE, which is also swept here by the rotor blades B over the entire circulation, comes to lie above the wing arrangement W.
  • the distance between the legs of the legs can of course be made larger than shown here, but at the same time should be larger than the largest outer diameter of the bladed rotor R.
  • FIG. 9 shows a greatly simplified exemplary sketch of a rotor R bladed with rotor blades B, left side view and right front view, and clarifies the definition of rotor planes RP in the region of the longitudinal extent of the bladed rotor R along its axis of rotation RA, these rotor planes RP always normal to the rotation axis RA.
  • a first rotor plane RP normal to the axis of rotation RA of the bladed rotor R can be defined in front, which, seen in the direction of flight FR, thereby just touching the front tip of a leading rotor blade of the rotor R touchingly attached; a last rotor plane RP with respect to the longitudinal extent of the rotor R along its axis of rotation RA touches, just as seen in the flow direction, one end of a far backward in the direction of flow rotor blade B and this plane is of course perpendicular to the axis of rotation RA of the bladed rotor R.
  • Layers are formed within the longitudinal extent of the rotor along its axis of rotation RA (the corresponding area in which rotation planes RP can ever occur, is indicated by arrows between the two planes). Each of these planes is normal to the axis of rotation RA and reaches an extent in the area that exceeds the spatial extent of the aircraft.
  • the planes of rotation RP serve as geometric auxiliary planes to detect the potential impact area of a bladed rotor R on the aircraft.

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  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Mechanical Engineering (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

L'invention concerne un aéronef sur lequel au moins un rotor à pales (R) entoure en rotation le fuselage (F) pour produire une propulsion, et est disposé d'une manière donnée par rapport aux composants de l'aéronef, et en même temps aussi les turbines d'un ensemble moteur (E), lesquelles entraînent au moins un tel rotor (R) par le biais de transmissions (G), sont disposées sur l'aéronef d'une manière définie de telle sorte qu'il en résulte avantageusement, pour l'aéronef, un fonctionnement sûr, à rendement élevé et à faible bruit.
PCT/DE2014/000007 2013-01-10 2014-01-10 Aéronef à rendement élevé et à faible bruit WO2014108125A1 (fr)

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Application Number Priority Date Filing Date Title
DE112014000391.3T DE112014000391B4 (de) 2013-01-10 2014-01-10 Lärmarmes und hocheffizientes Flugzeug

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DE102013000315.6 2013-01-10
DE102013000315 2013-01-10
DE102013015364.6A DE102013015364A1 (de) 2013-09-10 2013-09-10 Angepasste Flugzeugkonfigurationen für die Energieeffiziente Open-Rotor Integration
DE102013015364.6 2013-09-10

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WO2014108125A1 true WO2014108125A1 (fr) 2014-07-17

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Cited By (8)

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EP2964530B1 (fr) * 2013-03-09 2019-10-30 Rolls-Royce Corporation Groupe motopropulseur d'avion
DE102018208297A1 (de) * 2018-05-25 2019-11-28 Rolls-Royce Deutschland Ltd & Co Kg Luftfahrzeug mit mindestens einem Düsenantrieb
WO2021023712A1 (fr) * 2019-08-05 2021-02-11 Conseil Et Technique Aeronef
US11486472B2 (en) 2020-04-16 2022-11-01 United Technologies Advanced Projects Inc. Gear sytems with variable speed drive
US11535392B2 (en) 2019-03-18 2022-12-27 Pratt & Whitney Canada Corp. Architectures for hybrid-electric propulsion
US11628942B2 (en) 2019-03-01 2023-04-18 Pratt & Whitney Canada Corp. Torque ripple control for an aircraft power train
US11697505B2 (en) 2019-03-01 2023-07-11 Pratt & Whitney Canada Corp. Distributed propulsion configurations for aircraft having mixed drive systems
US11732639B2 (en) 2019-03-01 2023-08-22 Pratt & Whitney Canada Corp. Mechanical disconnects for parallel power lanes in hybrid electric propulsion systems

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GB2468917A (en) * 2009-03-26 2010-09-29 Circino Flight Systems Ltd Aircraft propulsion unit having two sets of contra-rotating, ducted propellers
WO2012114047A1 (fr) * 2011-02-25 2012-08-30 Airbus Operations Aéronef à impact environnemental réduit

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DE1884174U (de) 1961-03-22 1963-12-05 Boelkow Entwicklungen Kg Triebfluegelflugzeug.
GB2468917A (en) * 2009-03-26 2010-09-29 Circino Flight Systems Ltd Aircraft propulsion unit having two sets of contra-rotating, ducted propellers
WO2012114047A1 (fr) * 2011-02-25 2012-08-30 Airbus Operations Aéronef à impact environnemental réduit

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2964530B1 (fr) * 2013-03-09 2019-10-30 Rolls-Royce Corporation Groupe motopropulseur d'avion
DE102018208297A1 (de) * 2018-05-25 2019-11-28 Rolls-Royce Deutschland Ltd & Co Kg Luftfahrzeug mit mindestens einem Düsenantrieb
US11628942B2 (en) 2019-03-01 2023-04-18 Pratt & Whitney Canada Corp. Torque ripple control for an aircraft power train
US11697505B2 (en) 2019-03-01 2023-07-11 Pratt & Whitney Canada Corp. Distributed propulsion configurations for aircraft having mixed drive systems
US11732639B2 (en) 2019-03-01 2023-08-22 Pratt & Whitney Canada Corp. Mechanical disconnects for parallel power lanes in hybrid electric propulsion systems
US11535392B2 (en) 2019-03-18 2022-12-27 Pratt & Whitney Canada Corp. Architectures for hybrid-electric propulsion
WO2021023712A1 (fr) * 2019-08-05 2021-02-11 Conseil Et Technique Aeronef
US11486472B2 (en) 2020-04-16 2022-11-01 United Technologies Advanced Projects Inc. Gear sytems with variable speed drive

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DE112014000391B4 (de) 2021-06-17

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