Landing method and system for air vehicles

Patent metadata

Field	Value
Publication number	US-10423168-B2
Application number	US-201415108711-A
Country	US
Kind code	B2
Filing date	Dec 9, 2014
Priority date	Oct 7, 2014
Publication date	Sep 24, 2019
Grant date	Sep 24, 2019

Abstract

Methods are provided for operating an air vehicle, the air vehicle including fixed wings configured to provide mild stall characteristics including a post-stall regime, and a propulsion system capable of generating a controllable thrust, the thrust being variable at least between an idle thrust and a maximum thrust. During a landing maneuver, the air vehicle is caused to attain an angle of attack corresponding to said post-stall regime, and during the landing maneuver, there is concurrently generated a thrust level of said thrust greater than said idle thrust to provide a thrust vector having a thrust lift force component at landing. Corresponding control systems are also provided, and air vehicles including such control systems are also provided.

First claim

The invention claimed is:

1. A method for operating an air vehicle, the air vehicle comprising fixed wings configured to provide mild stall characteristics, wherein said wings exhibit a relationship of lift coefficient with angle of attack—including a post-stall regime comprising a plateau of said lift coefficients at a range of post-stall angles of attack greater than a stall angle of attack, and a propulsion system capable of generating a controllable thrust, the thrust being variable at least between an idle thrust and a maximum thrust, the method comprising:
during a landing maneuver, causing the air vehicle to attain an angle of attack corresponding to said plateau of said lift coefficients of said post-stall regime; and
during said landing maneuver, concurrently generating a thrust level of said thrust greater than said idle thrust to provide a thrust vector having a thrust lift force component at landing.

2. The method according to claim 1 , further comprising at least one of the following:
operating the air vehicle during said landing maneuver at airspeeds below a corresponding speed safety margin, while the air vehicle is airborne;
operating the air vehicle during said landing maneuver at airspeeds not exceeding the corresponding stall airspeed of the air vehicle, while the air vehicle is airborne; or
operating the air vehicle during said landing maneuver at airspeeds between 90% and 95% of the corresponding stall airspeed of the air vehicle, while the air vehicle is airborne.

3. The method according to claim 1 , wherein said lift coefficient in said plateau of said lift coefficients of said post-stall regime remains at 2.5+/−0.1 for a range of angle of attack between 10° and 20°.

4. The method according to claim 1 , wherein at least one of the following:
the air vehicle includes a UAV;
the air vehicle includes a UAV, and wherein the air vehicle has a weight between 50 kg to about 500 kg; or
the air vehicle has a weight between about 50 kg to about 500 kg.

5. The method according to claim 1 , further comprising, at least during said landing maneuver, generating significantly elevated drag levels as compared with nominal drag levels prior to said landing maneuver.

6. The method according to claim 1 , wherein at least one of the following:
said angle of attack is greater than 10°; or
said angle of attack is between 10° and 20°.

7. The method according to claim 1 , wherein at least one of the following:
said thrust lift force component is between 10% and 50% of said thrust level; or
said thrust level is between 10% and 50% of said maximum thrust.

8. The method according to claim 1 , wherein at least one of the following:
in said landing maneuver, the air vehicle is in a downward glide path defined by a glide angle; or
in said landing maneuver, the air vehicle is in a downward glide path defined by a glide angle, and, wherein said glide angle is in the range of between −3° to 15°.

9. The method according to claim 8 , wherein at least one of the following:
said thrust vector is at a pitch angle (θ) to the horizontal, wherein said pitch angle (θ) is related to said angle of attack (α) and said glide angle (γ) by the expression: α=θ−γ;
said thrust vector is at a pitch angle (θ) to the horizontal, wherein said pitch angle (θ) is related to said angle of attack (α) and said glide angle (γ) by the expression: α=θ−γ, and, wherein said thrust lift force component (TL) is related to said thrust level (T) and said thrust vector pitch angle (θ) by the expression: TL=T*sin θ;
said thrust vector is at a pitch angle (θ) to the horizontal, wherein said pitch angle (θ) is related to said angle of attack (α) and said glide angle (γ) by the expression: α=θ−γ, and, wherein said thrust vector has a forward thrust component (TH) at landing, wherein said forward thrust component (TH) is related to said thrust level (T) and said thrust vector pitch angle (θ) by the expression TH=T*cos θ; or
said thrust vector is at a pitch angle (θ) to the horizontal, wherein said pitch angle (θ) is related to said angle of attack (α) and said glide angle (γ) by the expression: α=θ−γ, and, wherein said thrust lift force component (TL) is related to said thrust level (T) and said thrust vector pitch angle (θ) by the expression: TL=T*sin θ, and, wherein said thrust vector has a forward thrust component (TH) at landing, wherein said forward thrust component (TH) is related to said thrust level (T) and said thrust vector pitch angle (θ) by the expression TH=T*cos θ.

10. The method according to claim 8 , wherein said glide angle (γ) is related to said thrust level (T), said elevated drag (D), and a lift force (L) by the expression: γ=(D−T)/L.

11. The method according to claim 1 , wherein said air vehicle is configured for providing said elevated drag levels while not significantly degrading aerodynamic lift provided by said wings during said landing maneuver.

12. The method according to claim 11 , wherein said fixed wings are lift generating wings having at least some aerofoil cross-sections thereof comprising a two element aerofoil, having an aerofoil chord, a primary element having a first leading edge and a first trailing edge, a secondary element having a second leading edge and a second trailing edge, a gap between the primary element and the secondary element, and an axial overlap between the first trailing edge and the second leading edge, the secondary element being deflectable with respect to the primary element about a fixed hinge point by a flap deflection angle, the secondary element being configured to operate in airbrake mode when deflected by a respective said flap deflection angle corresponding to a design airbrake deflection angle wherein to generate an airbrake drag, wherein at least for said design airbrake deflection angle said axial overlap is numerically greater than −0.5% of the aerofoil chord.

13. The method according to claim 12 , wherein at least one of the following:
said design airbrake deflection angle is greater than 40°;
said design airbrake deflection angle is greater than 45°;
said design airbrake deflection angle is greater than 55°.

14. The method according to claim 12 , wherein the primary element is configured for providing high lift mild stall characteristics, and wherein the aerofoil is configured to generate said airbrake drag while concurrently retaining said mild stall characteristics.

15. The method according to claim 12 , wherein at least for said design airbrake deflection angle said axial overlap is between −0.5% and +4% of the aerofoil chord.

16. The method according to claim 12 , wherein said axial overlap provides a smooth variation of aerofoil maximum lift coefficient with said flap deflection angle for a range of said flap deflection angles at least ±10° from said design airbrake deflection angle.

17. The method according to claim 12 , wherein at said axial overlap, the aerofoil maximum lift coefficient is maintained constant within 0.1 with said flap deflection angle for a range of said flap deflection angles at least ±10° from said design airbrake deflection angle.

18. The method according to claim 12 , wherein a value for said axial overlap is chosen to maximize the maximum lift coefficient obtained for the aerofoil at said design airbrake deflection angle.

19. The method according to claim 12 , wherein said secondary element is in the form of a slotted flap.

20. The method according to claim 12 , wherein the secondary element is deflectable with respect to the primary element abo

Assignees

• Israel Aerospace Ind Ltd

Inventors

• Shepshelovich Michael
• Abramov Danny

Classifications

• B64U70/60

  Take-off or landing of UAVs from a runway using their own power · CPC title

• B64D45/04

  Landing aids; Safety measures to prevent collision with earth's surface · CPC title

• B64C9/20

  by multiple flaps · CPC title

• B64C9/323

  associated with wings · CPC title

• B64C3/14

  Aerofoil profile · CPC title