RELATED APPLICATION

The instant application is related to commonly-owned, co-pending application entitled, “Unmanned Aerial Vehicle with Counter-Rotating Ducted Rotors And Shrouded Pusher-Prop”.

TECHNICAL FIELD

The present invention relates to unmanned aerial vehicles (UAVs). In particular to a high-speed vertical takeoff and landing (VTOL) UAV having a shrouded pair of coaxial, couner-rotating, ducted rotors for providing zero and low speed horizontal and vertical thrust, and a wing and shrouded pusher prop for providing forward translational lift and thrust in high-speed flight.

BACKGROUND OF THE INVENTION

There has been a recent increased emphasis on the use of unmanned aerial vehicles for performing various activities in both civilian and military situations where the use of manned flight vehicles is not appropriate and/or feasible. Such missions include surveillance, reconnaissance, target acquisition and/or designation, data acquisition, communications relay, decoy, jamming, harassment, ordnance delivery, or supply flights. This increased emphasis on the role of UAV's in today's (and tomorrow's) society has led to many advancements in both airframe design and propulsion systems.

There are generally three types of UAV configurations under current development, a fixed-wing type configuration (a fuselage with wings and horizontally mounted engines for translational flight), helicopter type configuration (a fuselage with a rotor mounted above which provides lift and thrust) and ducted type configuration (a fuselage with a ducted rotor system which provides translational flight, as well as vertical take-off and landing capabilities). A wingtype UAV provides several benefits over a helicopter or ducted type UAV. First, and foremost, is the ability of a winged UAV to travel at considerably greater speeds and for longer distances than a conventional helicopter or ducted type UAV. Also, a winged UAV can typically carry a larger mission payload and/or fuel supply than a helicopter or ducted type UAV. As such, fixed-wing UAVs are generally better suited than helicopter or ducted type UAVs for certain mission profiles involving endurance, distance, higher speed and load capability.

Winged UAVs, however, have deficiencies that limit their utility. For example, since winged UAVs require forward motion to maintain lift and therefore are not capable of hovering over a fixed spatial point. As a result, winged UAVs are not very good at delivering ordinance or laser designating targets. Also, winged UAVs cannot take-off and land vertically. Instead, winged UAVs require elaborate launch and retrieval equipment.

Helicopter UAVs can hover over a fixed spatial point and takeoff and land vertically but have limitations when operating in confined areas due to the exposed rotors rotating above the fuselage. Also, helicopter UAVs tend to have a high center-of-gravity (CG) and therefore have limited ability when landing on sloped surfaces or pitching ship decks. A high CG aircraft tends to roll over when landing on steep slopes.

The ability of ducted rotor-type UAVs to take-off and land vertically, combined with their ability to hover for extended periods of time over a point and operate in confined areas off steep slopes, make a ducted type UAVs ideally suited for real time tactical reconnaissance, target acquisition, surveillance, and ordnance delivery missions for front line tactical units.

Ducted-type UAVs, such as the CYPHER® unmanned aerial vehicle developed by Sikorsky Aircraft Corporation and generally disclosed in U.S. Pat. No. 5,152,478, includes a toroidal fuselage shrouding co-axial, counter-rotating rotors. The rotors are designed to provide the thrust necessary for both vertical and translational flights. As shown in

FIG. 1A

, aircraft vertical motion of the UAV is provided by maintaining the vehicle fuselage substantially horizontal so that the thrust (downwash) of the rotors provides the necessary lift for the aircraft. When fore-aft or lateral movement of the aircraft is desired, the aircraft fuselage must be “nosed-down” as shown in

FIG. 1B

in order to generate a horizontal thrust component.

As discussed above, ducted-type UAVs have a relatively slow speed as compared to winged UAVs. One reason fot this is that most ducted-type UAV's do not have a separate translational propulsive system. As such, the rotor system must provide both vertical and translational thrust, thus, requiring the full potential of the rotor system to be split.

Another problem associated with a toroidal UAV relates to drag. Referring back to

FIG. 1A

, if the aircraft were to fly in the forward direction (i. e., to the left in the figure) while oriented horizontally, the airflow passing over the nose N of the aircraft would impact the inner rear wall D

W

of the duct. This generates considerable drag on the aircraft. To reduce the drag on the aircraft, it is oriented as shown in FIG.

1

B. This orientation of the aircraft causes the airflow to pass through the rotor system, reducing airflow contact with the duct wall D

W

.

Another problem associated with conventional ducted-type UAVs is the tendency of the aircraft to experience a nose-up pitching moment. That is, the airflow over the airframe and through the rotor system produces a moment about the aircraft's center of gravity which causes the nose of the aircraft to pitch upward. There have been several attempts made to counter-act this nose-up pitching moment. U.S. Pat. No. 5,152,478 discloses a UAV rotor system wherein cyclic pitch is used to counter-act the nose-up pitching moment during forward translational flight. Although this solution does eliminate the nose-up pitching moment, it also requires a considerable amount of power and does not eliminate the drag on the duct wall.

Another possible option to counter-act the nose-up pitching moment is to optimize the toroidal fuselage airfoil profile. The utilization of an optimized toroidal fuselage airfoil profile to counteract the nose-up pitching tendency of UAVs is disclosed in U.S. Pat. No. 5,150,857. This solution requires that the outer aerodynamic surface of the toroidal fuselage be optimized to provide an asymmetrical toroidal fuselage pressure distribution that produces high lift forces during forward translational flight modes. The high lift forces reduce the required lift provided by the rotor assembly, thereby reducing the undesirable nose-up pitching moment. A reduction in required power is effected by the decreased requirement for rotor lift and the reduced need for superimposed cyclic pitch (moment trim).

While the incorporation of a toroidal fuselage having an optimized outer aerodynamic surface represents a viable option to help counteract the fuselage-induced nose-up pitching moments, this option incurs a manufacturing penalty and may have an adverse effect on higher speed flight characteristics.

A further solution to reducing the nose-up pitching moment in a ducted-type UAV is disclosed in U.S. Pat. No. 5,419,513 wherein ancillary wing structures are incorporated onto the aircraft to counter-act the nose-up pitching moment. More particularly, the ancillary wing structures have an aerodynamic configuration that generates lifting forces to supplement the lifting forces generated by the rotor assembly and the toroidal fuselage.

Although the incorporation of ancillary wings on the aircraft does help counter the nose-up pitching moment caused by translational flight, the aircraft translational thrust is still limited by the amount of horizontal thrust component that can be generated by the rotor system.

A need, therefore, exists for an improved rotor-type UAV which provides increased propulsive capabilities and reduced drag during forward translational flight.

SUMMARY OF THE INVENTION

An object of the present invention is to provide a method for reducing the nose-up pitching moment that occurs during forward flight of a unmanned aerial vehicle.

This and other objects and advantages of the invention are provided by the novel method of controlling airflow over an unmanned aerial vehicle according to the present invention. The unmanned aerial vehicle includes counter-rotating rotor assemblies that are mounted within a duct. Each rotor assembly includes a plurality of rotor blades. The method involves the step of adjusting the rotor blades to have substantially zero pitch. The rotors assemblies are then rotated to produce a virtual plane across the duct. The virtual plane is operative for substantially deflecting the air passing over the fuselage away from the duct.

In one embodiment of the invention, the method involves the further step of obstructing at least a portion of the bottom of the duct to inhibit air that is flowing across the bottom of the duct from passing into the duct.

In a further embodiment of the invention, a negative pressure is generated aft of the duct by increasing the amount of air flowing past the duct on the lower surface of the fuselage. The negative pressure creates a suction force on the air passing over the top of the duct.

As discussed above, ducted rotor aircraft encounter significant nose up pitching moments during forward flight. The present invention counteracts this nose up pitching moment by controlling several aspects of the aircraft. First, the airflow over and under the duct is controlled to minimize the airflow contacting the back wall of the duct. During high speed flight, additional counteracting forces may be needed to prevent the nose-up moments. The present invention addresses this need by creating lifting forces aft of the rotor assemblies, these lifting forces create counter-acting moments that further reduce the nose-up pitching moments. In one embodiment of the invention, the lift forces are created by controlling flaperons on wings mounted to the sides of the aircraft. Additionally, directional lanes located aft of a pusher prop can be controlled to deflect propeller trust downward creating additional lift.

It has also been determined through wind tunnel testing that locating the pusher prop assembly aft of the duct results in reduced drag on the aircraft. More particularly, the pusher prop assembly has been found to draw the turbulent air over the duct that would otherwise flow into the duct and cause drag in high-speed flight. The shroud around the propeller augments this beneficial affect.

Past attempts at high speed ducted rotor aircraft have used rigid retractable covers to conceal the rotor system and reduce drag. Such covers are relatively heavy and complex. In the present invention, during high-speed flight, the aircraft is in a near level body attitude with the rotors operating with nearly a flat pitch. This creates the virtual covers described above which impedes airflow into the duct. By blocking the flow and forcing it to flow over and under the aircraft, drag is reduced without the weight and complexity of rigid covers.

The virtual covers reduce the drag on the aircraft by about one half as much as two rigid covers. As discussed above, in one embodiment of the invention the bottom the bottom of the duct may be obstructed to prevent the entrance of air. The obstructing is provided by one or more deflectors mounted across a portion of the duct to control airflow into the duct from the bottom of the fuselage during forward flight. These deflectors are passive and require no actuation. They are opened and closed automatically based on the airflow through and over the duct. By being passive they save considerable complexity and weight over actively actuated covers. The combination of flat pitch rotors and passive flow deflectors on the bottom of the duct reduces the drag component on the aircraft between the open rotor and completely covered duct by about 80%.

The foregoing and other features and advantages of the present invention will become more apparent in light of the following detailed description of the preferred embodiments thereof, as illustrated in the accompanying figures. As will be realized, the invention is capable of modifications in various respects, all without departing from the invention. Accordingly, the drawings and the description are to be regarded as illustrative in nature, and not as restrictive.

BRIEF DESCRIPTION OF THE DRAWINGS

For the purpose of illustrating the invention, the drawings show a form of the invention which is presently preferred. However, it should be understood that this invention is not limited to the precise arrangements and instrumentalities shown in the drawings.

FIG. 1A

is a schematic representation of a prior rotor-type unmanned aerial vehicle during hover.

FIG. 1B

is a schematic representation of a prior rotor-type unmanned aerial vehicle during forward translational flight.

FIG. 2

is a perspective view of an unmanned aerial vehicle according to the present invention.

FIG. 3

is a perspective view of one embodiment of the unmanned aerial vehicle in

FIG. 2

illustrating many of the internal components.

FIG. 4A

is a schematic representation of the unmanned aerial vehicle according to the present invention during hover.

FIG. 4B

is a schematic representation of the unmanned aerial vehicle according to the present invention during forward translational flight.

FIG. 5

is a schematic representation of a preferred embodiment of the unmanned aerial vehicle according to the present invention.

FIG. 6

is a perspective view of one embodiment of the unmanned aerial vehicle which illustrates the wings detached from the fuselage.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Referring now to the drawings wherein like reference characters identify corresponding or similar elements throughout the several views,

FIG. 2

illustrates a perspective view of one embodiment of an unmanned aerial vehicle (UAV)

10

according to the present invention. The UAV

10

includes a fuselage

12

with a toroidal portion

14

having a generally hemicylindrical aerodynamic profile.

A rotor assembly

16

is mounted within a duct

18

that extends substantially vertically through the fuselage

12

. A plurality of struts

20

extend between the fuselage and the rotor assembly

16

. In the preferred embodiment illustrated in the figures, there are three struts

20

mounted in the fuselage

12

. The support struts

20

are rigidly attached to the rotor assembly

16

and the fuselage

12

, and are operative for supporting the rotor assembly

16

in a fixed co-axial relation with respect to the duct

18

. The support struts

20

also provide structural rigidity to the aircraft duct

18

to prevent flight and ground loads from distorting the fuselage

12

.

In order to minimize weight, the support struts

20

are preferably formed as hollow structures that can be used as conduits for interconnecting operating elements of the UAV

10

. For example, the engine drive shaft (identified by the numeral

36

is

FIG. 3

) is routed through one of the support struts

20

. In addition, the electrical wiring for various operating components in the aircraft

10

may be routed through the other support struts

20

.

The fuselage

12

and the plurality of support struts

20

can be made from various types of high tensile strength materials, such as composite and metal. The fuselage

12

includes a plurality of accessible internal bays (shown in

FIG. 3

) for housing and/or storing aircraft flight and mission components. For example, referring to

FIG. 3

which illustrates one embodiment of the UAV

10

according to the present invention, the compartments may be used to house a powerplant subsystem

22

, including fuel tanks

24

, flight control equipment

26

including data transmitters and receivers, a power conditioning unit

28

, and sensors

66

, such as a forward looking infrared radar (FLIR) sensor. In an alternate UAV configuration (shown in FIG.

5

), engines are mounted within the bays and the avionics equipment is mounted in the aft fuselage. The various components are preferably positioned within the compartments to balance the aircraft about the center of the rotor assembly.

The structural configuration of the UAV

10

in one embodiment is described in more detail in U.S. Pats. Nos. 5,152,478, 5,277,380, and 5,419,513, which are incorporated herein by reference in their entirety.

The powerplant subsystem

22

includes one or more fuel tanks

24

, an engine

30

, a drive train assembly

32

, and a cooling fan

34

. The fuel tanks

24

are disposed within appropriate internal bays, preferably in opposed equipment bays to maintain a relatively constant center of gravity for the UAV

10

during flight operations.

A drive train assembly

32

is operative for transferring power developed by the engine

30

to the rotor assembly

16

by means of a drive shaft

36

. The drive train assembly

32

may be of the type disclosed in commonly-owned, U.S. Pat. No. 5,226,350, which is incorporated herein by reference in its entirety.

In an alternate, and more preferred embodiment of the UAV

100

shown in

FIG. 5

, there are two engines mounted within the internal bays. Each engine

30

is preferably mounted forward of a fuel tank

24

. Mufflers

31

can be mounted adjacent to each engine

30

to reduce engine noise levels. The engines

30

are each engaged with the rotor system

16

by means of a drive shaft

36

that extends through a support strut

20

. The shafts

36

are preferably constructed of thin-wall aluminum tube for high strength and low weight. Flex couplings (not shown) are used to attach the shaft

36

to the engine

30

and the rotor assembly

16

to permit angular misalignment while transmitting torque.

The two engine embodiment shown in

FIG. 5

allows the aircraft to continue to operate in the event of an engine failure. An overrunning clutch is incorporated into the system to automatically disengage an inoperative engine from the system. The engines

30

are preferably mounted to the airframe with four shock mounts. In light of the location of the engine mounts in the front of the airframe, the mounts can also be used to support the landing gear.

A cross-feed line

102

is preferably attached to each fuel tank to allow fuel to migrate between the tanks

24

. This permits that aircraft to maintain balance under normal flight conditions. It also allows both engines to be fed in the case of one tank

24

running dry.

In the embodiment of the invention shown in

FIG. 5

, a separate cooling fan is not needed to cool the engines since air can be channeled into the compartment via an inlet on the airframe

12

.

The rotor assembly

16

includes a pair of multi-bladed, counter-rotating rotors

38

,

40

, coaxially aligned with the centerline of the duct

18

, and a coaxial transmission subassembly (not shown). The coaxial transmission subassembly may be of the type disclosed in commonly-owned, co-pending U.S. Pat. No. 5,226,350. The rotors

38

,

40

are aerodynamically “shrouded” by the fuselage

12

.

The counter-rotating rotors

38

,

40

are preferably of the rigid rotor type (as opposed to articulated rotors) to reduce the complexity and weight of the rotor assembly

16

. Each counter-rotating rotor

38

,

40

preferably includes a plurality of blade assemblies, which may be of the type disclosed in commonly-owned, co-pending U.S. Pat. No. 5,364,230, which is incorporated herein by reference in its entirety. In this type of rotor design, blade pitch changes induced in the counter-rotating rotors

38

,

40

, i.e., cyclic and/or collective pitch inputs, can be utilized to generate lift, pitch, roll, and yaw control of the UAV

10

. In the embodiment of the UAV

100

shown in

FIG. 5

, there are four blade assemblies on each counter-rotating rotor

38

,

40

.

In one preferred embodiment of the invention, the blades on the blade assemblies are rigid hingeless metal blades. The hingeless metal design is aerodynamically efficient, less costly than a composite blade, allows for additional control authority, and is less susceptible to foreign object damage. The blades preferably have a constant chord, are linearly twisted by about 25 degrees, and utilize a low Reynolds number airfoil. Alternatively, the blade can be made from wood with a metal leading edge and wrapped with a thin layer of fiberglass. The blades can also be tapered in chord and airfoil thickness from root to tip depending upon mission application. The blade tip speed of the rotor system is preferably about 62500 fps. This blade tip speed results in a rotor system that has low noise and reduced structural weight.

As will be discussed in more detail below, the pitch of the rotors can be adjusted to control the airflow over the shroud

12

and into the rotor assembly

16

. Such control of the airflow can create a lifting component on the shroud

12

that augments the lift provided by the counter-rotating rotors

38

,

40

.

In one embodiment of the UAV

10

, wings

42

are attached to and extend laterally outward from the sides of the aircraft fuselage

12

. More particularly, the fuselage

12

, in addition to its toroidal portion

14

that surrounds a portion of the duct

18

, includes an aft fuselage portion

44

or empennage (also referred to herein as a fairing). The aft fuselage

44

tapers as it extends rearwardly from the toroidal portion

14

of the shroud

12

. The wings

42

are preferably attached to the aft fuselage

44

at approximately its midplane.

The aerodynamic configuration of the wings

42

of the UAV

10

is selected to provide high lifting forces and a large nose-down pitching moment in forward translational flight. In addition, the wings

42

have a low aspect ratio so that the UAV

10

is maintained at a reasonable size. The wings

42

preferably taper outward from the fuselage

12

. Each wing

42

preferably has a symmetric or low camber airfoil section. Those skilled in the art would readily appreciate the diverse wing arrangements that can be incorporated into a UAV according to the present invention. Wing arrangements are dictated by the desired aircraft performance.

Preferably, at least a portion of the wings

42

is pivotally mounted so that the angle of attack of the wings

42

can be varied depending on the aircraft's mission. For example, in one embodiment of the invention shown in

FIG. 3

, the entire wing

42

could be pivotally mounted to the airframe. The wings

42

would then be angled so as to provide additional lift when the aircraft is in a loiter mode of operation. Conversely, the wings

42

would be highly angled if the mission requires a high degree of maneuverability. To control pivoting of the wings

42

, an actuator

45

is mounted within the fuselage

12

near each wing

42

. The actuator

45

is engaged with a drive rod

46

that is attached to the wing

42

. The actuator

45

is preferably an electrical servo that is controlled by signals from an onboard computer that regulates the amount that the actuator rotates the drive rod

46

about its longitudinal axis. The rotation of the drive rod

46

changes the angle of attack of the wing

42

. As is shown in the illustrated embodiment, there are two actuators

45

, one for controlling each wing

42

. This allows the wings

42

to be independently pivoted. It is also contemplated that a single actuator can be used to pivot both wings if independent pivoting is not desired.

In a more preferred embodiment of the invention shown in

FIG. 5

, only a portion of the wing

42

is pivotable. In this embodiment, each wing

42

includes a fixed stub portion

42

F

and a pivotal portion

42

P

. The pivotal portion

42

P

preferably includes a flaperon hingedly mounted to the trailing edge of the wing

42

. An actuator

45

mounted within the fixed portion

42

F

controls the pivoting of the pivotal portion

42

P

. The flaperon

42

P

preferably has a deflection range from about 45 degrees trailing edge down to about 15 degrees trailing edge up. As shown in

FIG. 5

, it is preferable that the pivotal portion

42

P

also includes a pivotal tip end or cap

42

T

which functions as an aerodynamic and mass balance to facilitate flap actuation. The pivoting tip end reduces control loads on the actuator

45

by approximately 55%, thereby increasing actuator life and reliability. Those skilled in the art would readily appreciate the diverse pivotal wing arrangements that can be incorporated into a UAV according to the present invention.

Referring to

FIG. 6

, the wings

42

are preferably removably attached to the fuselage

12

. This permits operational personnel to detach the wings

42

from the fuselage

12

during storage or transit. The wings

42

are then reattached to the fuselage

12

prior to use. The wings are mounted using conventional bolts or quick disconnect couplings. Since the actuator

45

is mounted within the fixed portion

42

F

of the wing in the embodiment of the invention shown in

FIG. 5

, only electrical control lines need to connect from the UAV fuselage

12

to the wing

42

.

It is contemplated that wings with different geometries and aerodynamic characteristics can be attached to the airframe

12

depending on the specific aircraft mission. Also the fuselage

44

has a straight or flat interface surface on the left and right sides which forms a mating surface for the wing. This smooth flat surface simplifies wing attachment and removal.

In order to provide translational thrust, the UAV

10

includes a usher prop

50

mounted to the rear of the aircraft

10

. More particularly, the pusher prop (or propeller)

50

is mounted to the rear of the aircraft with its rotational axis oriented substantially horizontal. The pusher prop

50

includes a plurality of blades

56

, which in the illustrated embodiment there are two blades

56

with a diameter of about

20

inches. The propeller

50

is mounted to a drive shaft

58

which, in turn, is engaged with a power unit. In the preferred embodiments, the power unit which drives the propeller

50

is the engine

30

that is used to power the rotor assembly

16

. As shown in

FIG. 5

, the propeller drive shaft

58

is engaged with the engines

30

though the rotor systems gearbox. A flexible coupling is incorporated between the gearbox and the drive shaft

58

to allow for shaft misaligments. The pusher prop

50

is designed to provide the aircraft's forward thrust in high-speed flight.

The propeller

50

operates at about 7000 RPM and preferably is of a hingeless rigid design. The blades

56

are mounted so that the collective pitch of the blades can be controlled. In order to improve aircraft agility in low speed flight, the propeller is designed with reverse pitch capability. The propeller

50

in the preferred embodiment is designed with a tip speed of about 637 fps but it could be operated at any other suitable propeller tip speed.

A prop shroud

54

is formed on the aft fuselage

44

around the pusher prop

50

and preferably has a cylindrical or fiustoconical shape. The cross-sectional shape of the shroud

54

is preferably configured as an airfoil to provide the shroud

54

with some lift. The lift on the shroud

54

creates a moment about the aircraft's center of gravity that stabilizes the aircraft in the pitch and yaw directions. The shroud

54

also protects the propellers from environmental contact, increases the propeller efficiency and reduces the aircraft's acoustic signature.

Mounted on the shroud

54

aft of the pusher prop

50

are one or more directional vanes

60

. The vanes

60

are configured to direct the exhaust from the pusher prop

50

in a particular direction. Preferably, the vanes

60

are pivotally mounted to the shroud

54

to permit the exhausted air to be channeled in a controllable manner. For example, during transition it is desirable to angle the vanes

60

so as to channel the air downwardly, thereby augmenting the vertical lift provided by the rotor assembly

16

. In high-speed flight, the vanes

60

function as elevators on the aircraft to provide pitch control for the vehicle.

Referring to

FIGS. 2

,

4

A and

4

B, the aft fuselage

44

preferably has a contour which assists in channeling the air into the pusher prop

50

. In the illustrated embodiment, the top surface of the aft fuselage

44

angles downwardly toward the shroud

54

. Similarly, the lower surface of the aft fuselage

44

angles upward toward the shroud

54

. This results in a channeling of the airflow into the pusher prop

50

.

As discussed above with respect to

FIG. 1B

, during forward flight of a conventional toroidal fuselage ducted-type UAV, the airflow into the duct impacts the wall of the duct D

W

causing considerable drag on the aircraft and increases the nose up pitching moment experienced by the fuselage. For the near level body high speed flight orientation of this invention, air also enters from the bottom of the duct and impacts the wall of the duct D

W

and causes the same large drag increase. In one embodiment of the present invention, the UAV

10

reduces the drag caused during forward flight by controlling the flow of air into the duct

18

.

Referring to

FIG. 4A

, the UAV

10

is shown in its vertical takeoff and hover mode with the airflow

70

funneling downward through the rotor assembly

16

as is conventional in the prior UAV designs.

FIG. 4B

illustrates the UAV

10

according to the present invention in high-speed flight. As shown, the airflow

70

′ that passes over top of the nose of the aircraft tends to angle downward into the rotor assembly

16

. To prevent the air from entering the duct

18

, the rotors

38

,

40

are adjusted so as to have approximately zero pitch, i.e., substantially flat pitch. At near zero pitch, the rotational speed of the rotors

38

,

40

create a virtual plane

62

across the top and bottom of the duct

18

. This virtual plane

62

, which is coplanar with the rotors

38

,

40

, deflects the air passing over and under the vehicle away from the duct

18

. As a result, the passing air does not impact the rear wall of the duct D

W

. Instead, the air is channeled into the pusher prop

50

. The pitch of the rotors

38

,

40

is changed slightly both collectively and cyclically as the air vehicle's flight speed changes to minimize drag. The benefits of the virtual plane

62

are primarily present at flight speeds where the aircraft is supported entirely by the wings

42

and fuselage

12

.

To further reduce aircraft drag the present invention incorporates one or more deflectors

64

on the lower surface of the aircraft. The deflectors prevent airflow from entering the duct

18

from the lower surface of the UAV

10

. Referring to

FIGS. 4A and 4B

, the UAV

10

is shown with deflectors

64

mounted to the lower surface of the fuselage

12

below the rotor assembly

16

. The deflectors

64

are attached to the aircraft via hinges and are free to pivot.

When in the hover mode, the downwash from the rotors causes the deflectors

64

to open as shown in FIG.

4

A. In forward flight, the natural high pressure on the bottom of the aircraft causes the deflectors

64

to close off the duct, thereby preventing airflow

70

″ from entering into the duct from the lower surface. As such, the airflow

70

″ across the bottom of the UAV

10

passes the duct

18

and is channeled into the pusher prop

50

as shown in FIG.

4

B.

A 40% scale model of a UAV

10

incorporating many of the aspects of the invention described above was tested to assess the effect that rotor speed has on drag. The test results proved that by increasing the speed of the rotors

38

,

40

, the overall drag on the UAV

10

could be reduced. The UAV was also tested with the deflectors

64

. The incorporation of the deflectors

64

reduced drag on the aircraft by almost 80% over an unpowered open rotor configuration.

The low speed flight control and flying qualities of a UAV according to the present invention are similar to those described in U.S. Pat. No. 5,152,478. The high speed control and flying qualities of the UAV

10

, on the other hand, are similar to conventional fixed wing aircraft. In high speed flight, the pivotable flap portions on the wings act as ailerons for providing roll control.

Directional vanes in the shroud function as elevators. While a vertical vane is not shown within the shroud in the figures, it could readily be added to provide yaw control if mission requirements dictate its need.

In order to transition the UAV

10

from hover flight mode to forward flight mode, the rotor assemblies, flaperons and pusher prop must be controlled to blend low speed rotor control into high-speed wing and elevator control. The following is a discussion of one very efficient transition profile. The aircraft is capable of flying in transition mode anywhere from zero to 20 degrees of body attitude. To begin transition to fixed wing mode the UAV

10

is pitched about five degrees nose down. At about 5 kts the aircraft begins to rotate nose up. The aircraft reaches 10 degrees angle of attack upon reaching 20 kts. The aircraft remains at 10 degrees angle of attack out to 55 kts. As the aircraft accelerates from 20 kts to 55 kts, rotor thrust is gradually reduced as wing lift increases. This is described in U.S. Pat. No. 5,152,478 and U.S. Pat. No. 5,150,857, which are both incorporated herein in there entirety.

The UAV according to the present invention is capable of zero degree flaperon wing borne flight at 55 kts and will begin to nose down as speed is further increased. Above 55 kts the UAV flies like a typical fixed wing aircraft. Transition back to helicopter mode is accomplished by reversing the steps described above.

Although the invention has been described and illustrated with respect to the exemplary embodiments thereof, it should be understood by those skilled in the art that the foregoing and various other changes, omissions and additions may be made therein and thereto, without parting from the spirit and scope of the present invention.