Aircraft ground effect altimeter for autonomous landing control
First Claim
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1. A method for automatic aircraft landing control using ground effect to determine how close an aircraft is to the ground, comprising:
- (a) providing an electronic processor configured to communicate with a navigation unit on an aircraft, said aircraft configured with an electronic instrument cluster for takeoff and landing control, said aircraft selected from the group of aircraft consisting of fixed and rotary-winged air vehicles, said fixed wing air vehicles having a wingspan and said rotary-winged air vehicles having a rotary diameter;
(b) receiving an aircraft model (M) from said navigation unit of said aircraft;
(c) using said aircraft model (M) to compute aircraft modeled quantities, said aircraft modeled quantities including;
aircraft accelerations ({umlaut over (x)},ÿ
,{umlaut over (z)}), rotation velocities ({dot over (α
)},{dot over (β
)},{dot over (γ
)}), and altitude (h) as a function of aircraft surface control settings (C), thrust settings (T), and hypothesized ground altitude (g) at time (t), wherein (m) denotes said aircraft modeled quantities, wherein Pm=[{umlaut over (x)},ÿ
,{umlaut over (z)},{dot over (α
)},{dot over (β
)},{dot over (γ
)},h]m=M(g,C,T,t), wherein, •
Pm are state parameters corresponding to said aircraft modeled quantities;
(d) initializing said time (t) at (t=0);
(e) receiving an initial altitude (b) of said aircraft from a barometric altimeter on said aircraft, wherein initial aircraft velocities ({dot over (x)},{dot over (y)},ż
) and initial aircraft orientations (α
,β
,γ
) are from said navigation system;
(f) initializing said aircraft model (M) as M(b,{dot over (x)},{dot over (y)},ż
,α
,β
,γ
);
(g) receiving accelerations, rotation velocities, and altitude from the output of the accelerometers, gyros, and barometric altimeter valid for said time (t), where (α
) denotes the actual measurements of said aircraft, wherein Pα
=[{umlaut over (x)},ÿ
,{umlaut over (z)},{dot over (α
)},{dot over (β
)},{dot over (γ
)},h]α
, wherein Pα
are state parameters corresponding to said actual measurements of said aircraft;
(h) receiving the surface control (C) and thrust (T) settings, at time (t), from said instrument cluster;
(i) computing a best hypothesis for the ground altitude (g) with lowest error (e) searching over g in the range (g=h to h−
l), where (l) is the length of said wingspan when said aircraft is a fixed wing air vehicle or said rotary diameter when said aircraft is a rotary-winged air vehicle, and denoting (gmin) as the error minimizing hypothesized ground altitude, wherein said lowest error (e) is defined by
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Abstract
Embodiments of the invention use ground effect to determine how close an aircraft is to the ground. An electronic processor communicates with a navigation unit on an aircraft. The aircraft has an electronic instrument cluster for takeoff and landing control. An aircraft model is received from the navigation unit and used to compute aircraft modeled quantities. An initial altitude of the aircraft is received from a barometric altimeter. Initial aircraft velocities and initial aircraft orientations are from the navigation unit. The aircraft model is initialized. Accelerations, rotation, and altitude from the output of accelerometers, gyros, and the barometric altimeter are received. Surface and thrust control settings are received from the electronic instrument cluster. A best hypothesis for ground altitude is computed and the computed ground altitude having the lowest likelihood of error is reported.
11 Citations
10 Claims
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1. A method for automatic aircraft landing control using ground effect to determine how close an aircraft is to the ground, comprising:
-
(a) providing an electronic processor configured to communicate with a navigation unit on an aircraft, said aircraft configured with an electronic instrument cluster for takeoff and landing control, said aircraft selected from the group of aircraft consisting of fixed and rotary-winged air vehicles, said fixed wing air vehicles having a wingspan and said rotary-winged air vehicles having a rotary diameter; (b) receiving an aircraft model (M) from said navigation unit of said aircraft; (c) using said aircraft model (M) to compute aircraft modeled quantities, said aircraft modeled quantities including;
aircraft accelerations ({umlaut over (x)},ÿ
,{umlaut over (z)}), rotation velocities ({dot over (α
)},{dot over (β
)},{dot over (γ
)}), and altitude (h) as a function of aircraft surface control settings (C), thrust settings (T), and hypothesized ground altitude (g) at time (t), wherein (m) denotes said aircraft modeled quantities, wherein Pm=[{umlaut over (x)},ÿ
,{umlaut over (z)},{dot over (α
)},{dot over (β
)},{dot over (γ
)},h]m=M(g,C,T,t), wherein, •
Pm are state parameters corresponding to said aircraft modeled quantities;(d) initializing said time (t) at (t=0); (e) receiving an initial altitude (b) of said aircraft from a barometric altimeter on said aircraft, wherein initial aircraft velocities ({dot over (x)},{dot over (y)},ż
) and initial aircraft orientations (α
,β
,γ
) are from said navigation system;(f) initializing said aircraft model (M) as M(b,{dot over (x)},{dot over (y)},ż
,α
,β
,γ
);(g) receiving accelerations, rotation velocities, and altitude from the output of the accelerometers, gyros, and barometric altimeter valid for said time (t), where (α
) denotes the actual measurements of said aircraft, wherein Pα
=[{umlaut over (x)},ÿ
,{umlaut over (z)},{dot over (α
)},{dot over (β
)},{dot over (γ
)},h]α
, wherein Pα
are state parameters corresponding to said actual measurements of said aircraft;(h) receiving the surface control (C) and thrust (T) settings, at time (t), from said instrument cluster; (i) computing a best hypothesis for the ground altitude (g) with lowest error (e) searching over g in the range (g=h to h−
l), where (l) is the length of said wingspan when said aircraft is a fixed wing air vehicle or said rotary diameter when said aircraft is a rotary-winged air vehicle, and denoting (gmin) as the error minimizing hypothesized ground altitude, wherein said lowest error (e) is defined by - View Dependent Claims (2, 3, 4, 5)
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6. A non-transitory computer readable medium having stored thereon a plurality of computer executable instructions for determining how close an aircraft configured with a navigation system is to the ground, that when executed by a computer including a GPU, causes the computer to:
-
(a) receive an aircraft model (M) from a navigation system of an aircraft, said aircraft having an electronic instrument cluster for takeoff and landing control, said aircraft selected from the group of aircraft consisting of fixed and rotary-winged air vehicles, said fixed wing air vehicles having a wingspan and said rotary-winged air vehicles having a rotary diameter; (b) use said aircraft model (M) to compute aircraft modeled quantities, said aircraft modeled quantities including;
aircraft accelerations ({umlaut over (x)},ÿ
,{umlaut over (z)}), rotation velocities ({dot over (α
)},{dot over (β
)},{dot over (γ
)}), and altitude (h) as a function of aircraft surface control settings (C), thrust settings (T), and hypothesized ground altitude (g) at time (t), wherein (m) denotes said aircraft modeled quantities, wherein Pm=[{umlaut over (x)},ÿ
,{umlaut over (z)},{dot over (α
)},{dot over (β
)},{dot over (γ
)},h]m=M(g,C,T,t), wherein Pm are state parameters corresponding to said aircraft modeled quantities;(c) initialize said time (t) at (t=0); (d) receive an initial altitude (b) of said aircraft from a barometric altimeter on said aircraft, wherein initial aircraft velocities ({dot over (x)},{dot over (y)},ż
) and initial aircraft orientations (α
,β
,γ
) from an navigation system;(e) initialize said aircraft model (M) as M (b,{dot over (x)},{dot over (y)},ż
,α
,β
,γ
);(f) receive accelerations, rotation velocities, and altitude from the output of the accelerometers, gyros, and barometric altimeter valid for said time (t), where (α
) denotes the actual measurements, wherein Pα
=[{umlaut over (x)},ÿ
,{umlaut over (z)},{dot over (α
)},{dot over (β
)},{dot over (γ
)},h]α
, wherein Pα
are state parameters corresponding to said actual measurements of said aircraft;(g) receive the surface control (C) and thrust (T) settings, at time (t), from said instrument cluster; (h) compute a best hypothesis for the ground altitude (g) with lowest error (e) searching over g in the range (g=h to h−
l), where (l) is the length of said wingspan when said aircraft is a fixed wing air vehicle or said rotary diameter when said aircraft is a rotary-winged air vehicle, and denoting (gmin) as the error minimizing hypothesized ground altitude, wherein said lowest error (e) is defined by - View Dependent Claims (7, 8, 9, 10)
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Current Assigneethe united states of america as represented by the secretary of the navy
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Original Assigneethe united states of america as represented by the secretary of the navy
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InventorsHilde, Jeffrey J.
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Primary Examiner(s)Shaawat, Mussa A
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Assistant Examiner(s)Kim, Kyung
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Application NumberUS14/314,264Time in Patent Office664 DaysField of SearchNoneUS Class Current1/1CPC Class CodesB64D 45/04 Landing aids; Safety measur...G01C 23/00 Combined instruments indica...G01C 5/005 altimeters for aircraft G01...G05D 1/042 specially adapted for aircraftG05D 1/0676 specially adapted for landing
