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Title:
AIRCRAFT FOR TRANSPORT AND DELIVERY OF PAYLOADS
Document Type and Number:
WIPO Patent Application WO/2017/042291
Kind Code:
A1
Abstract:
The invention relates to a vertical take-off and vertical landing aircraft (1,1b) configured for flying in a first flight mode (200) and a second flight mode (201), wherein the aircraft (1,1b) comprises: -a front section (2) comprising two fixedly attached front wings (20), -a means for propulsion (6), wherein the means for propulsion (6) is configured to provide propulsion, when the aircraft (1,1b) flies in the first flight mode (200), and wherein the means for propulsion (6) is configured to provide lift, when the aircraft (1,1b) flies in the second flight mode (201), - a rear section (3) comprising two rear wings (30), wherein the front wings (20) and the rear wings (30) each are configured to provide lift, when the aircraft (1,1b) flies in the first flight mode (200), - a longitudinal axis (5) that extends from the rear section (3) to the front section (2), wherein the longitudinal axis (5) points in a normal flight direction (202) with the front section (2) facing forward, when the aircraft (1,1b) flies in the first flight mode (200), wherein oa rear point (32) on the longitudinal axis (5) is provided by the normal projection of the aerodynamic center (31) of the rear wings (30) onto the longitudinal axis (5), oa front point (22) on the longitudinal axis (5) is provided by the normal projection of the aerodynamic center (21) of the front wings (20) onto the longitudinal axis (5), and oa center point (CP) on the longitudinal axis (5) is provided by the normal projection of the center of gravity (CG) of the aircraft (1,1b) onto the longitudinal axis (5), and wherein the aircraft (1,1b) is configured such that when the aircraft (1,1b) is in a take-off position (401) or landing position (400), the longitudinal axis (5) encloses and angle of more than 45° with the ground with the front section (2) facing upwards, wherein the center point (CP) is between the rear point (32) and the front point (22) and is spaced apart from the rear point (32) more than 15% of the distance between the rear point (32) and the front point (22) and wherein the means for propulsion (6) is arranged at the front section (2).

Inventors:
WEISS JONAS (CH)
Application Number:
EP2016/071223
Publication Date:
March 16, 2017
Filing Date:
September 08, 2016
Export Citation:
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Assignee:
SWISS AEROBOTICS AG (CH)
International Classes:
B64C29/02; B64C39/02; B64C39/08
Domestic Patent References:
WO2015028627A12015-03-05
Foreign References:
US20140339354A12014-11-20
US4387866A1983-06-14
FR2675114A11992-10-16
FR2993245A12014-01-17
US20130206921A12013-08-15
Other References:
None
Attorney, Agent or Firm:
SCHULZ, Ben Jesko (Großbeerenstraße 71, Berlin, DE)
Download PDF:
Claims:
Claims

A vertical take-off and vertical landing aircraft (1 ,1 b) configured for flying in a first flight mode (200) and a second flight mode (201 ), wherein the aircraft (1 ,1 b) comprises:

a front section (2) comprising two fixedly attached front wings (20), a means for propulsion (6), wherein the means for propulsion (6) is configured to provide propulsion, when the aircraft (1 ,1 b) flies in the first flight mode (200), and wherein the means for propulsion (6) is configured to provide lift, when the aircraft (1 ,1 b) flies in the second flight mode (201 ), a rear section (3) comprising two rear wings (30), wherein the front wings (20) and the rear wings (30) are each configured to provide lift, when the aircraft (1 ,1 b) flies in the first flight mode (200),

a longitudinal axis (5) that extends from the rear section (3) to the front section (2), wherein the longitudinal axis (5) points in a normal flight direction (202) with the front section (2) facing forward, when the aircraft (1 ,1 b) flies in the first flight mode (200), wherein

o a rear point (32) on the longitudinal axis (5) is provided by the

normal projection of the aerodynamic center (31 ) of the rear wings (30) onto the longitudinal axis (5),

o a front point (22) on the longitudinal axis (5) is provided by the normal projection of the aerodynamic center (21 ) of the front wings (20) onto the longitudinal axis (5), and

o a center point (CP) on the longitudinal axis (5) is provided by the normal projection of the center of gravity (CG) of the aircraft (1.1 b) onto the longitudinal axis (5),

and wherein the aircraft (1 ,1 b) is configured such that when the aircraft (1 .1 b) is in a take-off position (401 ) or landing position (400), the longitudinal axis (5) encloses an angle of more than 45° with the ground with the front section (2) facing upwards,

characterized in that

the center point (CP) is located between the rear point (32) and the front point (22), and wherein the center point (CP) is spaced apart from the rear point (32) more than 15% of the distance between the rear point (32) and the front point (22), and wherein the means for propulsion (6) is arranged at the front section (2). Aircraft according to claim 1 , characterized in that the center point (CP) is spaced apart from the rear point (32) more than 20%, particularly more than 30%, preferably more than 40% and more preferably 50% of the distance between the rear point (32) and the front point (22).

Aircraft according to claim 1 or 2, characterized in that the means for propulsion (6) is arranged on each front wing (20).

Aircraft according to one of the preceding claims, characterized in that the means for propulsion (6) comprises a propeller (60, 60a).

Aircraft according to one of the preceding claims, characterized in that the means for propulsion (6) comprises two propellers (60, 60a), wherein each propeller (60, 60a) is arranged on one of the front wings (20).

Aircraft according to claim 4 or 5, characterized in that the propeller (60, 60a) diameter of said propeller (60, 60a) or of said propellers (60, 60a) is greater than 70% of the semi-wingspan (24) of the respective front wing (20).

Aircraft according to one of the preceding claims, characterized in that the first flight mode (200) is a fixed-wing flight mode, wherein the fixed-wing flight mode is characterized in that the lift for the aircraft (1.1 b) is generated by the front wings (20) and the rear wings (30), when the aircraft (1 ,1 b) flies in the fixed-wing flight mode, and wherein the lift is particularly caused by the forward airspeed and the shape of the front and rear wings (20, 30), and/or wherein the second flight mode (201 ) is a rotary-flight mode, wherein in the rotary-flight mode the lift is generated by rotating blades (61 ) of a propeller (60, 60a) that revolve around a shaft.

Aircraft according to one of the preceding claims, characterized in that the means for propulsion (6) is tiltable around at least one axis with respect to the aircraft (1 ,1 b) such that a thrust vector of the means for propulsion (6) is changed, wherein if the means for propulsion (6) is a propeller (60, 60a), the plane that comprises the rotatable blades (61 ) of the propeller (60, 60a) is tiltable around at least one axis.

9. Aircraft according to one of the preceding claims, characterized in that the propeller (60, 60a) comprises means for controlling the angle of attack of the blades (61 ) comprised by the propeller (60, 60a), wherein the means for controlling the angle of attack is particularly configured to reciprocally change the angle of attack within a full turn of the propeller (60, 60a), wherein said means for controlling the angle of attack particularly comprises a helicopter swash-plate. 0. Aircraft according to claim 5 to 9, characterized in that the aircraft (1 ,1 b) comprises flight control surfaces wherein, when the aircraft (1 .1 b) flies in the first flight mode (200) or in the second flight mode (201 ), the flight control surfaces (23, 23i, 23o, 33) are arranged in the downwash of the rotating propeller (60) or the rotating propellers (60), particularly such that control authority is provided independently of the flight speed of the aircraft (1.1 b).

1 1 . Aircraft according to one of the preceding claims, characterized in that the aircraft (1 ,1 b) is configured such that the aircraft's flight attitude is

controllable, particularly when the aircraft (1 ,1 b) flies in the second flight mode (201 ), through a differential thrust provided by the means for propulsion (6), wherein the differential thrust is provided by a variable power-setting of the means of propulsion (6), preferably by a variable pitch of the propeller (60, 60a) or particularly by a differential or split orientation by the front wing (20) control surfaces (23, 23i, 23o). 12. Aircraft according to one of the preceding claims, characterized in that the aircraft (1.1 b) comprises a payioad segment (4) wherein the payioad segment (4) is removably arranged between the front section (2) and the rear section (3), wherein the aircraft (1 , 1 b) and the payioad segment (4) comprise means for attaching and releasing the payioad segment (4) to or from the aircraft (1 , 1 b), wherein the payioad segment (4) particularly comprises the center of gravity (CG) and/or the center point (CP) of the aircraft (1 ,1 b) and wherein the center of gravity of the payioad segment (4) is nearby or on the center of gravity (CG) of the aircraft (1 ,1 b).

13. Aircraft according to one of the preceding claims, characterized in that the rear wings (30) are configured to serve as a landing support (71 ) when the aircraft (1 ,1 b) is in the take-off (401 ) or landing position (400) and wherein preferably also a single rear fin (34) or a plurality of rear fins (34) of the aircraft (1.1 b) are configured to serve as a landing support (71 ).

14. Aircraft according to one of the preceding claims, characterized in that the rear section (3) comprises an extendable landing support system (70) that is extendable between a compact position and an extended position, wherein said landing support system (70) is arranged such at the rear section (2) of the aircraft (1 ,1 b) that a tip-over inclination angle (80) of the aircraft (1 ,1 b) is increased when the landing support system (70) is in the extended position, when the aircraft (1 ,1 b) is arranged in a take-off position (401 ) or a landing position (400), and wherein particularly the landing support system (70) is configured to adopt the compact position when the aircraft (1 ,1 b) flies in the first flight mode (200), wherein the landing support system (70) is particularly arranged at the rear wings (20) and/or the rear fins (34).

15. System comprising an aircraft according to one of claims 1 to 14 and a

plurality of removably arrangabie payioad segments (4), wherein the payioad segments (4) are particularly configured such that they can be connected to each other so as to form a new payioad segment that is particularly removably arrangabie between the front section (2) and the rear section (3) of the aircraft (1 ,1 b) and wherein particularly each payioad segment (4) itself is also removably arrangabie between the front section (2) and the rear section (3), wherein the aircraft (1 , 1 b) and the payioad segment (4) or the new payioad segment comprise means for attaching and releasing the payioad segment (4) to or from the aircraft (1 , 1 b), wherein the payioad segment (4) or the new payioad segment particularly comprises the center of gravity (CG) and/or the center point (CP) of the aircraft (1 ,1 b) and wherein the center of gravity of the payioad segment (4) or the new payioad segment is nearby or on the center of gravity (CG) of the aircraft (1 ,1 b).

Description:
Aircraft for transport and delivery of payloads Specrjjcajion

The invention relates to a vertical take-off and vertical landing aircraft according to claim 1 and a system comprising such an aircraft according to claim 15.

Vertical take-off and vertical landing (VTOL) aircrafts are adapted for landing in regions where no runways can be provided.

The most popular kind of VTOL aircrafts is the helicopter of the class of rotary wing aircrafts.

The main limitation of helicopters is their low air speed, low aerodynamic efficiency and a resulting short mission range.

While fixed-wing aircrafts resolve the issue of low speed and short mission ranges, they generally need a landing strip and a runway in order to land or take-off.

For this reason, hybrid aircrafts have been designed that combine the advantages of rotary-wing and fixed-wing aircrafts, particularly aircrafts that combine the two flight modes. These kinds of VTOL aircrafts comprise for example the class of so-called tail-sitter aircrafts that take off and land on their rear and then tilt horizontally for forward flight.

Tail-sitters however are comparably complicated to land for a pilot sitting in such a tail-sitter aircraft, as the pilot typically faces upwards with his back pointing towards the tail. Configurations with a single propulsion system for both flight modes are known.

While tail-sitters were considered impractical solutions for many decades, computational power for flight robotics has evolved to a level that particularly unmanned aerial vehicles (UAV) are comparably easy and reliably to control.

Increasing efforts are made to adapt UAVs for payload delivery in regions where landing sites only provide limited space - for example in form of multi-rotor aircrafts, such as quadro-copters, hexa-copters and the like. The advantage of using such multi-rotors is that almost arbitrary payloads can be attached for example below such an aircraft without impacting its flight attitude considerably. However, as stated above, helicopters, multi-rotor aircrafts and aircrafts configured only for rotary-wing flight have limited mission range and offer only low airspeed and are thus not well suited for a fast delivery of payloads and a sufficiently large mission range. Attaching arbitrary payloads to other kinds of VTOLs on the other hand will affect their flight attitude particularly in fixed wing flight mode considerably as with varying payloads the center of gravity shifts along the longitudinal axis of the fuselage, altering the flight Attitude particularly the pitch (pitch motion) attitude of the aircraft. Therefore, the problem underlying the present invention is to provide an aircraft allowing the fast and long-range transport and delivery of payloads to regions where landing and take-off space is limited.

This problem is solved by an aircraft having the features of claim 1 . Preferred embodiments are stated in the sub claims.

According to claim 1 , an aircraft for vertical take-off and vertical landing configured for flying in a first flight mode, particularly a fixed-wing flight mode, and a second flight mode, particularly a rotary-flight mode, comprises a front section comprising two particularly fixedly attached front wings, a means for propulsion, wherein the means for propulsion is configured to provide propulsion, when the aircraft flies in the first flight mode, and wherein the means for propulsion is configured to provide lift, when the aircraft flies in the second flight mode,

a rear section comprising two particularly fixedly attached rear wings, wherein the front wings and the rear wings each are configured to provide particularly positive lift, when the aircraft flies in the first flight mode, a longitudinal axis that extends from the rear section to the front section, wherein the longitudinal axis points in a normal flying direction with the front section facing forward and the rear section facing backward, when the aircraft flies in the first flight mode, wherein

o a rear point on the longitudinal axis is provided by the orthogonal projection of the aerodynamic center of the rear wing onto the longitudinal axis,

o a front point on the longitudinal axis is provided by the orthogonal projection of the aerodynamic center of the front wing onto the longitudinal axis, and

o a center point on the longitudinal axis is provided by the

orthogonal projection of the center of gravity of the aircraft onto the longitudinal axis,

and wherein the aircraft is configured such that when the aircraft is in a take- off or landing position, or when the aircraft flies in the second flight mode, the longitudinal axis points away from the ground enclosing an angle of greater than 45° with the ground and with the front section facing upwards and particularly with the rear wings closer to the ground than the front wings, wherein the center point is located between the rear point and the front point, and is spaced apart from the rear point more than 15% of the distance between the rear point and the front point and, wherein the means for propulsion is arranged at the front section.

As the center point is arranged between the front point and the rear point, the weight and thus the lift in the first flight mode of the aircraft is distributed accordingly on the front wings and rear wings. This is particularly advantageous as the configuration of the aircraft comprising front wings and rear wings provides increased flight stability almost independent of the position of the center of gravity.

Such a configuration is also called a canard or tandem configuration of wings. In canard designs known form the state of the art, the center of gravity is located closer to the aerodynamic center point of the rear wing. Therefore these designs are not well suited for varying centers of gravity caused for example by varying payloads, as the trim capabilities of such canard designs are not sufficiently strong to provide stable operations.

The combination of the means for propulsion that is arranged at the front section of the aircraft and the tandem wing design of the aircraft advantageously combines results in an aircraft design that is well-suited for vertical take-off and vertical landing as well as for fixed wing flight.

Furthermore, the aircraft according to the invention advantageously combines the advantages of tail-sitter VTOL aircrafts with the advantages regarding the position of the center of gravity of payloads provided by helicopters and other rotary wing aircrafts, particularly through the tandem configuration of the wings. As the aircraft is arranged at an angle greater than 45° with respect to the ground when in landing or take-off position, it is possible to vertically take-off. In this context angles within the range of 80 - 1 10 are preferred, wherein theoretically and practically an angle of 90 is ideal.

The aircraft according to the invention is therefore advantageously suited for transport and delivery of payloads to regions that exhibit only limited landing and take-off space.

The front section is preferably not overlapping with the rear section of the aircraft. Furthermore the front section and the rear section of the aircraft are particularly sections of a fuselage of the aircraft. The distance between the front section and the rear section can be variably designed and is preferably alterable.

A means for propulsion particularly comprises an electric or a combustion engine.

The longitudinal axis extends lengthwise through the fuselage of the aircraft from the nose to the tail of the aircraft. Angular movement of the aircraft around the longitudinal axis is known as roll of the aircraft.

Furthermore, the longitudinal axis preferably passes through the center of gravity, such that the center point and the center of gravity are identical within close bounds.

According to another embodiment of the invention the center point is spaced apart from the rear point more than 20%, particularly more than 30%, preferably more than 40% and more preferably 50% of the distance between the rear point and the front point.

The exact position of the center of gravity influences the wing areas (also called the surface area of the wings) of the front wings and the rear wings. For example, if the center of gravity lies at 50% of the distance between rear point and front point, the area of the front wings and the rear wings will be almost the same. The closer the center of gravity is located to the rear of the aircraft, comparably the larger the wing area of the rear wings should be.

This embodiment of the invention is particularly robust with regard to varying payloads that are particularly arranged between the rear and front section of the aircraft. Given a set of flight-parameters (speed, thrust, etc.), the flight attitude of the aircraft remains almost unaffected by altering payloads particularly when the center point is in the middle of the front point and the rear point, i.e. the center point coincides with the aerodynamic center of the aircraft. This also holds true for varying absolute separation distances between the rear and the front-point, i.e. for different length and shapes of payload segments.

According to another embodiment of the invention, the means for propulsion is arranged on each front wing. This configuration is particularly advantageous as the means for propulsion provides control over the flight direction, for example through control surfaces in the downwash of the propulsion that control the pitch, roll and yaw motion and through differential thrust of the propulsion means for the yaw motion of the aircraft.

Furthermore, this configuration is advantageous in that with two propulsion units increased reliability in the first flight mode can be achieved. Another advantage of this configuration is that the propeller downwash does not hit the fuselage and a potential payload segment, thus aerodynamic, particularly propulsion efficiency is not affected by different sizes and shapes of the payload segment

According to another embodiment of the invention, the surface area of the front wings and the surface area of the rear wings have the same effective, aerodynamic size with a tolerance of up to 30%. The advantage of nearly identical wing surface areas is that both wing pairs - the front wings and the rear wings - can provide enough lift and control authority such that the aircraft remains operational almost independently of where the center point or center of gravity along the longitudinal axis is located between the front and rear point.

The surface area of a wing is particularly the surface area that is used for lift in first flight mode.

Alternatively to the surface areas of the wings, the front wings and the rear wings are configured to provide almost the same amount of lift when the aircraft flies in the first flight position.

According to another embodiment of the invention, the means for propulsion comprises a propeller.

The propeller is preferably configured as a proprotor. A proprotor is a propeller that is configured for providing propulsion for fixed-wing mode and to provide lift in rotary wing mode.

In a variant of the invention the means for propulsion comprises two propellers, wherein each propeller is arranged on one of the front wings.

This configuration advantageously provides control authority over the aircraft particularly with respect the generation of differential thrust for a yaw motion of the aircraft.

According to another embodiment of the invention the propeller diameter of particularly each of the propellers is greater than 70% of half the wingspan of the front wing. This configuration of an aircraft allows for big propellers that can provide enough lift in order to allow for controlled take-off and landing.

According to another embodiment of the invention, each propeller is arranged on the respective front wing at a position less than 50% of the semi-wing span of the front wings and wherein each propeller comprises a diameter that is greater than 70%, particularly greater than 100%, of the semi-wing span of the front wings and wherein each propeller encloses an angle with the longitudinal axis and wherein the propellers are arranged in and configured for intermeshing rotor operation. Providing intermeshing propellers allows for greater lift power while maintaining a compact aircraft design.

According to a preferred embodiment of the invention, the first flight mode is a fixed- wing flight mode, wherein the fixed-wing flight mode is characterized in that the lift for the aircraft is generated by the front wings and the rear wings, when the aircraft flies in the fixed-wing flight mode, and wherein the lift is particularly caused by the forward airspeed and the surface area and shape of the wings.

Along the same line of thinking, another preferred embodiment of the invention is characterized in that the second flight mode is a rotary flight mode, wherein in the rotary-flight mode the lift is generated by rotating propeller or proprotor blades that revolve on a shaft.

According to another embodiment of the invention, the means for propulsion is rotatable around at least one axis with respect to the aircraft such that a thrust vector of the means for propulsion is changed, wherein if the means for propulsion is a propeller, the plane that comprises the rotatable blades of the propeller is rotatable around at least one axis. The thrust-vector points in the direction of the thrust generated by the means for propulsion, wherein if the means for propulsion comprises two propellers, each propeller has its own thrust vector.

According to another embodiment of the invention the propeller comprises means for controlling the angle of attack of the blades comprised by the propeller, wherein the means for controlling the angle of attack is particularly configured to reciprocally change the angle of attack within a full turn of the propeller, wherein said means for controlling the angle of attack particularly comprises a swash-plate, conceptually similar to the ones used for helicopters.

According to another embodiment of the invention, the aircraft is configured such that the aircraft's flight attitude is controllable, particularly when the aircraft flies in the second flight mode, through a differential thrust provided by the means for propulsion, wherein the differential thrust is provided by an adjustable power-setting of the means of propulsion, preferably by a variable pitch of the propeller or control surfaces that are configured to act as spoilers on the propeller downwash.

According to another aspect of the invention, the aircraft comprises aircraft stabilizers, wherein the stabilizers are configured such that the aircraft maintains its directional flight stability even if one of the propellers is not providing thrust.

According to another embodiment of the invention, the aircraft comprises flight control surfaces wherein, when the aircraft flies in the first or second flight mode, the flight control surfaces are arranged in the downwash of the rotating propeller or the rotating propellers particularly such that control authority is provided independently of the flight speed the aircraft.

According to another embodiment of the invention the aircraft has a thrust-to-weight ratio greater one when the aircrafts flies in the first and/or second flight mode. Particularly providing a maximum thrust to weight ratio of up to 1 .5 in the second flight mode, will allow operating the aircraft under much broader and adverse operating/wind and payioad conditions.

According to another aspect of the invention the aircraft comprises a payioad segment, wherein the payioad segment is arranged between the front section and the rear section, wherein the payioad segment particularly comprises the center of gravity and/or the center point of the aircraft, wherein the center of gravity of the payioad segment is nearby or on the center of gravity of the aircraft, wherein more preferably the orthogonal projection of the center of gravity of the payioad segment on the longitudinal axis of the aircraft is on the center point of the aircraft. "Nearby" in this context means that the center of gravity of the payioad segment is spaced apart from the center of gravity of the airplane preferably less than 15% of the distance between the rear point and the front point.

This embodiment is advantageous as trimming of the aircraft when flying in first flight mode is kept at a minimum as regardless of the weight of the payioad in the payioad segment, the center of gravity of the aircraft does not change considerably.

According to another aspect of the invention the rear wings of the aircraft are configured to serve as a landing support when the aircraft is in the take-off or landing position and wherein also particularly a single one or a plurality of rear fins of the aircraft are configured to serve as the landing support.

Rear-fins are one kind of stabilizers for aircrafts. A landing support for the aircraft is particularly characterized in that it is capable of carrying the weight of the aircraft and dynamic forces, when the aircraft is in the landing or take-off position. Thus a landing support can be understood as the landing gear of the aircraft.

In a variation of the invention the rear section comprises an extendable landing support system that is extendable between a compact position and an extended position, wherein said landing support system is arranged such at the rear section of the aircraft that a tip-over inclination angle of the aircraft is increased when the landing support system is in the extended position and when the aircraft is arranged in a take-off or landing position, and wherein particularly the landing support system adopts the compact position when the aircraft flies in the first flight mode, wherein the landing support system is particularly arranged at the rear wings and/or the rear fin.

The increased tip-over tolerance for such kinds of aircrafts is particularly advantageous, as it allows for a higher tolerance with regard to the landing site or take-off site requirements, such as its inclination, as well as for landing more securely for example in windy weather conditions.

According to another embodiment of the invention the aircraft is an unmanned aerial vehicle. An unmanned aerial vehicle is particularly remote controlled or is capable of fully autonomous flight.

In another embodiment of the invention the surface area of the front wings is smaller than the surface area of the rear wings.

The problem according to the invention is furthermore solved by a system comprising an aircraft according to the invention and a plurality of removably arrangable payload segments, wherein the payload segment are configured particularly such that they are interconnectable such as to form a new payload segment that is removably arrangable between the front section and the rear section to the aircraft, particularly repeatedly. "Repeatedly" in this context means that each segment can be removed without destruction of any part of the airplane or the payload segment.

This particularly modular system of an aircraft with interchangeable payload segments is advantageously suited for delivery of various payloads as the payload segments can exhibit various shapes, lengths, weights and volumes. The system according to the invention is thus ideally suited to carry varying payloads with only a minimum need for re-trimming as long as the center of gravity of the aircraft remains between the limits set. Further features and advantages of the invention shall be described by means of a detailed description of embodiments with reference to the Figures, wherein it is shown in

Fig. 1 a top view the aircraft according to the invention;

Fig. 2 a front view of the aircraft according to the invention;

Fig- 3 a side view of the aircraft according to the invention;

Fig. 4 a perspective view of the aircraft according to the invention;

Fig- 5 a variation of the aircraft according to the invention;

Fig. 6 landing support system for increasing the tip-over angle;

Fig. 7 support points, and support polygons;

Fig. 8 a flight curve for take-off; Fig. 9 a flight curve for landing the aircraft;

Fig. 10 another flight curve for landing the aircraft;

Fig. 1 1 a schematic drawing of the front and rear wing and the associated distances between aerodynamic centers and various lift forces;

Fig. 12 a diagram showing the lift of the front and rear wing with regard to the angle of attack; and

Fig. 13 a schematic drawing showing the differential and split orientation of control surfaces. In Fig. 1 to Fig. 4 schematic views of an unmanned aircraft according to the invention are shown. Fig. 1 shows a top view, Fig. 2 shows a front view, Fig. 3 shows a side view and Fig. 4 shows a perspective view of the aircraft 1 .

The aircraft 1 is a tail-sitter aircraft with a tandem wing configuration. The term tail- sitter refers to the fact that in a landing position 400 and a take-off position 401 of the aircraft 1 , the aircraft 1 "sits" with a rear section 3 (the tail) on the ground.

The aircraft 1 comprises a fuselage 7 that comprises a front section 2 and a rear section 3, wherein between the front section 2 and the rear section 3 a payload segment 4 is arranged. At the front section 2 two front wings 20 are arranged. At the rear section 3 two rear wings 30 are arranged. Along the fuselage 7 a longitudinal axis 5 extends from the rear section 3 towards the front section 2. The longitudinal axis 5 is particularly the axis around which the aircraft 1 performs roll motions. A span-wise axis extends orthogonally to the longitudinal axis 5 and defines a pitch axis or lateral axis of the aircraft 1 . A third axis, the vertical axis or yaw axis, is the axis around which the aircraft 1 performs yaw motions.

An aerodynamic center 21 , 31 can be defined for the front wings 20 as well as the rear wings 31 , wherein the aerodynamic center 21 , 31 is the point on an airfoil where the aerodynamic (or pitching) moment remains almost constant with respect to the angle of attack.

A normal projection of the aerodynamic centers 21 of the front wings 20 onto the longitudinal axis 5 is defined as the front point 22 and the normal projection of the aerodynamic center 31 of the rear wings 30 onto the longitudinal axis 5 is defined as the rear point 32. The center of gravity CG of the aircraft 1 ideally lies also on the longitudinal axis 5 or in close proximity of it. Irrespectively of that, it is always possible to define a center point CP that lies on the longitudinal axis 5 and which is defined by the normal projection of the center of gravity CG onto the longitudinal axis 5.

The center point CP according to the invention lies between the rear point 32 and the front point 22 on the longitudinal axis 5. The design of the aircraft 1 according to this example is furthermore chosen such that the center point CP lies in the middle of the front point 22 and the rear point 32. Also, the center of gravity 41 of the payload segment coincides with the center of gravity CG of the aircraft 1 .

As the aircraft 1 is designed with a center point CP that is located such that the lift that has to be generated by the front wings 20 and the rear wings 30 is approximately the same in a first flight mode. The front wings 20 and the rear wings 30 are almost equal in size.

The first flight mode 200 is a so-called fixed-wing flight mode that is characterized by a lift that is generated by the airflow above and below fixed wings.

In contrast to the first flight mode, the aircraft 1 according to the invention is capable of a second flight mode 201 that is called rotary-wing flight mode. The rotary-wing flight mode is characterized by a lift that is generated by rotating wings, as for example by propellers.

For example, helicopters fly in rotary-flight mode, whereas jets fly in fixed-wing flight mode.

The aircraft 1 is capable of ta king-off in rotary-wing flight mode and subsequently transition to fixed-wing flight mode. The advantage of such a flight behavior is that only minimal starting space is required for takeoff and on the other hand, long flight distances can be covered comparably fast and efficient in fixed-wing flight mode. For the purpose of fixed-wing flight and rotary-wing flight the aircraft 1 comprises two propellers 6 that are attached one on each front wing 20. The propellers 60 are the means for propulsion 6 for both the fixed-wing flight mode and the rotary-wing flight mode. Furthermore the propellers 60 and their corresponding motors are arranged at the middle of each front wing 20. The rotor blades 61 of the propellers 60 are particularly large in order to provide a comparably large airflow over the surfaces of the front wings 20 and rear wings 30, wherein a focus lies on the fact that the airflow hits the control surfaces 23 of the front wings 20 and the rear wings 30 such that also in rotary-flight mode control can be maintained over the aircraft 1 even though the aircraft 1 is not moving at high velocity. In the present example, the diameter of each propeller 60 is approximately 70% of the semi-wing span 24 of the front wings 20. The controls in the second flight mode are particularly given by a pitch and roll motion wherein the yaw motion might be obtained by different thrust of the respective propeller 60 or by differential or split operation of the control surfaces 23.

The combination of the propellers 60 arranged at the front wings 20 and the center point CP lying close to the middle of the rear point 32 and front point 22 provides an aircraft 1 that is capable of vertical takeoff as well as vertical landing, and capable of flying in fixed wing mode without the need of elaborate trimming for each new payload as both - the front wings 20 and the rear wings 30 - carry essentially the same amount of weight, when flying in fixed-wing flight mode.

As the absolute weight and the location of the center point CP is important for the trimming in other aircraft designs, for example in canard designs where the front wings are comparably small with respect to the rear wings, the flight behavior of the present design is almost independent of the absolute weight of the payload.

The design of an aircraft 1 according to the invention provides a robust and reliable aircraft 1 that is capable of delivering varying payloads while maintaining flight attitude stability, particularly longitudinal stability, and air-speed at the same time.

Control surfaces 23, 33 are comprised by front wings 20 and rear wings 30. The front wings 20 and the rear wings 30 furthermore comprise stabilizers 35 at their respective end sections that serve as vertical and/or horizontal stabilizers.

Furthermore, the aircraft 1 comprises rear fins 34 that are considered stabilizers as well.

The rear wings 30 and the rear fins 34 have support means 71 attached to them. These support means 71 are configured such that the rear wings 30 and rear fins 34 are not damaged when the aircraft 1 is in the take-off or landing position and carry the full weight of the aircraft. Where the support means 71 touch the ground, they define support points 72 or lines that form a support polygon on the ground. This support polygon can be used to estimate the tip-over stability when the aircraft 1 is in the take-off 401 or landing position 400.

Regarding the effective, aerodynamic size of the front wings 20 and the rear wings 30 and the position of the center of CG gravity the following considerations can be made. The condition for natural pitch stability during flight is that the rear momentum Mrear must exhibit a stronger change with respect to a change of angle of attack a than the front momentum M fron t (see Fig. 10). The front and rear momentum can be estimated by the lift of wing L (L fron t , L rear respectively) times the distance d rea r, d fron t from the center of gravity CG or center point CP to the rear or front point respectively (see Fig. 1 1 ). The condition for pitch stability can be met for identical wing areas with the CG exactly in the middle between the two aerodynamic centers 21 , 32 by selecting the front wing section with a smaller L/a slope, but operating it at a higher angle of attack a, to create an intersection with the rear curve, i.e. provide identical lift as the rear wing but with smaller L/a slope. The intersection is the aircraft design operating point at a given airspeed.

Adjustable parameters are wing airfoils-sections, angle-of attack a and wing area.

Trade-offs for the adjustable parameters are: For high-efficiency, long wings are required (Prandtl elliptic lift distribution), which become limited by structural and weight considerations. One large wing is better than two smaller wings. Wing-to-wing interferences should be avoided and accounted for. If the aircraft 1 is "too stable", large control surfaces are required to control it, if the aircraft 1 is close to unstable (neutrally stable), large and fast rudders need to be used.

Fig. 5 shows a schematic view of another aircraft 1 b according to the invention, wherein said aircraft 1 b has propellers 60a that are arranged in the a pushing manner, as in contrast to the aircraft 1 shown in Fig. 1 to Fig. 4 where the propellers 60 are arranged in a pulling manner.

In Fig. 13 it is shown how flight stability in rotary wing mode (or hover mode) can be achieved by differential, resp. split operation of the control surfaces 23.

In general there are two ways, that can be used separately or combined in order to achieve or control yaw motions in rotary-flight mode. A yaw notion is either achieved through differential thrust by the means for propulsion as explained above - e.g. by a variable pitch or speed of the propellers - or through so-called dynamic spoilers that are configured to offset or cancel thrust by control surfaces 23 that in this context are also called deflecting surfaces 23, wherein said deflecting surfaces are preferably arranged at the front wings 20. Each front wing 20 comprises an inner 23i and an outer 23o deflecting surface, wherein the inner deflecting surface 23i is closer to the fuselage 7 than the outer deflecting surface 23o. In order to achieve or compensate a yaw motion of the aircraft 1 the inner and outer deflecting surfaces 23i, 23o, are deflected in opposite directions such that a drag is produced but no pitch motion (see Fig. 13 left panel). This configuration is called the differential orientation of the control surfaces.

Alternatively (see Fig. 13 right panel), at least one of the inner or outer deflecting surface 23 can be split into a first half 231 and second half 232, wherein the first and second half 231 , 232 are oriented in opposite directions such that a drag is produced but no pitch. This configuration is called the split orientation of the control surfaces.

The deflecting surfaces 23, 23i, 23o, 231 , 232 can be oriented by rotation parallel to the front wing 20 axis or aerodynamic front point 21.

Fig. 6 shows two aircrafts 1 according to the invention, in their take-off position 401 . The aircraft 1 shown on the right panel of Fig. 6 comprises an extendable support system 70, wherein the aircraft 1 shown in the left panel lacks thereof.

Fig. 6 particularly illustrates how a tip-over angle is increased by the extendable support system 70. The tip-over angle 80 is reached, when the center of gravity CG of the aircraft 1 lays outside the support polygon 74 which extents between the support points 72 or along the support lines. The maximum tip-over angle 80 of the two aircrafts 1 shown in Fig. 6 is different, as the corresponding support polygon 74 differs for the two aircrafts 1 . The extendable support system 70 is configured to increase the size of the support polygon 74 such that the tip-over angle 80 is increased. This is achieved by extensions 73 of the extendable support system 70 that are attached to the rear wings 30 and to the rear fins 34. These extensions 73 are particularly movable such they do not provide drag when the aircraft 1 is in first flight mode 200.

Fig. 7 shows a top view of the two aircrafts 1 shown in Fig. 6 with and without the extendable support system 70. It can be seen that the support points 72 for the aircraft 1 with the extendable support system 70 are spaced farther apart on the ground and thus provide an increased tip-over angle 80 for the aircraft 1 with the extendable support system 70.

Fig. 8 shows the take-off flight curve (dotted line) of the aircraft 1. It can be seen that the aircraft 1 transitions from rotary-wing flight mode to fixed-wing flight mode. The arrows on the dotted line indicate the normal flight direction 202.

Fig. 9 and Fig. 10 show two possible landing flight curves (dotted lines). It can be seen that the aircraft 1 switches from fixed-wing flight mode to rotary-wing flight mode in two different ways. The flight curve shown in Fig. 10 provides a shorter landing procedure wherein the landing flight curve according to Fig. 9 is a more secure way of landing the aircraft 1 according to the invention. The landing flight curve shown in Fig. 10 is particularly achievable when the propellers 60, 600 are tiltably arranged on the front wings 20. which adds control authority to the aircraft.

The arrows on the dotted line indicate the normal flight direction 202. The take-off 401 and landing position 400 of the aircraft 1 is orthogonal to the ground. I.e. the angle enclosed with the longitudinal axis 5 and the ground is 90 .