Source: http://www.google.com/patents/US8170730?dq=6,373,753
Timestamp: 2016-12-09 18:13:42
Document Index: 375434157

Matched Legal Cases: ['Application No. 11', 'Application No. 60', 'Application No. 200910174359', 'Application No. 2580272', 'Application No. 2580272', 'Application No. 2580272', 'Application No. 200580034374', 'Application No. 200910174359', 'Application No. 200910174359', 'Application No. 200580034374', 'Application No. 200580034374']

Patent US8170730 - Control system for automatic flight in windshear conditions - Google PatentsSearch Images Maps Play YouTube News Gmail Drive More »Sign inPatentsA flight control system is configured for controlling the flight of an aircraft through windshear conditions. The system has means for measuring values of selected flight performance states of the aircraft and a control system for operating flight control devices on the aircraft. A windshear detection...http://www.google.com/patents/US8170730?utm_source=gb-gplus-sharePatent US8170730 - Control system for automatic flight in windshear conditionsAdvanced Patent SearchTry the new Google Patents, with machine-classified Google Scholar results, and Japanese and South Korean patents.Publication numberUS8170730 B2Publication typeGrantApplication numberUS 13/192,522Publication dateMay 1, 2012Filing dateJul 28, 2011Priority dateOct 8, 2004Fee statusPaidAlso published asCA2580272A1, CA2580272C, CN100551777C, CN101068712A, CN101667036A, CN101667036B, DE05858232T1, EP1835835A2, EP1835835A4, EP1835835B1, US8000847, US20080046137, US20110282523, WO2007005045A2, WO2007005045A3, WO2007005045A9Publication number13192522, 192522, US 8170730 B2, US 8170730B2, US-B2-8170730, US8170730 B2, US8170730B2InventorsShyhpyng Jack ShueOriginal AssigneeTextron Innovations Inc.Export CitationBiBTeX, EndNote, RefManPatent Citations (15), Non-Patent Citations (21), Referenced by (6), Classifications (10), Legal Events (2) External Links: USPTO, USPTO Assignment, EspacenetControl system for automatic flight in windshear conditions
US 8170730 B2Abstract
1. A method of controlling the flight of an aircraft in windshear conditions, the method comprising:
measuring flight performance states of the aircraft while the aircraft flies under the control of a control system for operating flight control devices on the aircraft based at least in part on control inputs from a ground station;
calculating average gusts encountered by the aircraft with a computer-based system using a reverse system parameter identification function, a reverse spectral density function, and a time counter function;
calculating a plurality of covariance aircraft parameter through residual computation with the reverse system parameter identification function;
wherein the reverse spectral density function uses the plurality of covariance aircraft parameters to calculate average gust encountered by the aircraft;
comparing the average gusts to pre-determined values in a table; and
when the average gust exceeds a selected value in the table indicating windshear conditions, and based at least in part on communications with the ground station, automatically activating a recovery system for operating at least one of the flight control devices on the aircraft so as to minimize the effects of the windshear conditions on the flight of the aircraft.
automatically turning the heading of the aircraft into a horizontal gust, such that the gust becomes a headwind.
transmitting a warning signal to the ground station and activating the recovery system based at least in part on a state of communications with the ground station such that the recovery system is activated based at least in part on a determination that a predetermined amount of time has passed since control inputs were last received from the ground station.
minimizing of the effects of the windshear conditions includes ignoring control inputs from the ground station so as to keep the airspeed of the aircraft within a specified range.
5. The method of claim 1, wherein the aircraft is an unmanned aircraft.
This application is a divisional of U.S. Application No. 11/663,906, filed 26 Mar. 2007, now U.S. Pat. No. 8,000,847 titled “Control System for Automatic Flight in Windshear Conditions,” which claims the benefit of International PCT Application No. PCT/US05/36722, filed 11 Oct. 2005, titled “ Control System for Automatic Flight in Windshear Conditions,” which claims the benefit of Provisional Application No. 60/617,410, filed 8 Oct. 2004, titled “ Control System for Automatic Flight in Windshear Conditions,” all of which are hereby incorporated by reference for all purposes as if fully set forth herein.
The present invention relates generally to the field of flight control systems for aircraft and relates particularly to a system for automatic flight and recovery in microbursts.
The problem of controlling the flight and recovery of a UAV under microburst conditions on the flight path is shown in FIG. 1. As shown in the figure, UAV 11 is programmed to fly along path 13. A microburst area 15 in storm 17 produces a central downburst 19, which is directed outward into lateral outbursts 21, 23 when downburst 19 contacts the ground below storm 17. As UAV 11 passes through outburst 21, it may have extra lift generated by the upward wind coming from the ground or by a head wind, and this condition will be referred to as Condition 1. UAV 11 in this condition is “ballooning,” and the extra lift may cause the flight control system to reduce the flight control inputs due to this phenomenon. UAV 11 will then lose large amounts of lift when the direction and speed of winds change as UAV 11 flies into Condition 2, which is the central downburst 19.
The present invention is directed to a flight control system configured for automatically controlling the flight of an aircraft in a microburst. The system decreases the possibility of losing UAVs under microburst attack and increases the survivability rate, making UAVs having the system of the invention a more cost-effective system.
In order to make the UAV detect microburst conditions and intelligently recover from this impact, intelligent state flow technology is used in the system of the invention. In recent years, there has been considerable progress in the control of uncertain systems using H∞, robust control techniques. The H∞, robust technique has been shown to be the best method to account for the uncertainty in microburst models and aerodynamic coefficients in control law design.
In microburst, the lift of the aircraft is changed not only with respect to the maneuvers of the system, but also with respect to the related airspeed and air density. Therefore, the total lift of aircraft in the microburst contains large uncertain coefficient terms, and these coefficients are varying with respect to wind speed, wind direction, aerodynamics, angle of attack, etc. This is discussed below in greater detail.
This equation also contains some uncertain factors when calculated for conditions within a microburst.
a z x U 0 = α . - q - l x q . U 0 = h ¨ U 0 - l x q . U 0 , ( 3 ) where the parameter, lX, is the distance between the CG of the aircraft and the sensor measuring the acceleration.
The above lift, moment, and acceleration equations are considered as the most important terms affected by microburst. The reason for selecting these three variables is to minimize the sensitivity of trajectory changes in the flight path under microburst. Therefore, these terms are employed as the performance outputs of the system used for the H∞, robust control technique.
{dot over (X)} k lat =A k lat X lat +B k lat U lat Y k lat =C k lat X lat where k=0, 1, . . . j (4)
{dot over (X)} long =A k longt X long +B k long U long Y long =C k long X long where k=0, 1, . . . j (5)
where Xlat is (v, φ, ψ, p, r)T, Xlong is (u, w, θ, q)T, Ulat is (δped, δlat)T, and Ulong is (δlong, δcoll)T. Ak lat and Ak long are 5×5 and 4×4 matrices, respectively, Bk lat and Bk long are 5×2 and 4×2 matrices, and Ck lat and Ck long are 5×10 and 4×11 matrices. Note that k=0, 1, . . . j, are the number of models selected to compute the robust feedback control gain for the flight control system. The total numbers selected for the control law development are not limited and are dependent on system performance requirements. The above state space representation implies that all state variables are employed to provide H∞ robust feedback control laws. Because the system of the invention is primarily directed to a rotary-wing aircraft, the above equations of motions are chosen from the steady state value of aircraft speed between 60-100 kts with an altitude less than 1750 ft.
Unit direction gust
{ u = u + u g w = w + w g q = q + q g for longitudinal gust terms ( 6 ) { v = v + v g r = r + r g p = p + p g for lateral gust terms ( 7 ) Assume that ηlong=(ug wg qg) and ηlat=(vg rg pg). Applying (6) and (7) into equations (4) and (5) results in
{dot over (X)} lat =A k lat X lat +B k lat U lat +G lat V g lat Y lat =C k lat X lat +D lat V g lat Z lat =C k lat X lat where k=0, 1, . . . j (8)
{dot over (X)} long =A k long X long +B k long U long +G long V g long Y long =C k long X long +D long V g lat Z lat =C k lat X lat where k=0, 1, . . . j (9)
From equations (8) and (9), the difference between each model for k=0, 1, . . . j can be computed when aerodynamic parameters based on various airspeed sand densities are determined. After combining all differences, the bounds of these differences can be determined. Therefore, the whole group of linear systems on equation (8) and (9) can be rewritten as:
{dot over (X)} lat=(A lat +ΔA)X lat+(B lat +ΔB)U lat +G lat V g lat Y lat =C 1 lat X lat +V g lat Z lat=(C 2 lat +ΔC 2 lat)X lat (10)
{  Δ A j → j + 1 Lat  ≤ ξ Lat T Q ξ Lat ≤ Q _ Lat  Δ A j → j + 1 Long  ≤ ξ Long T Q ξ Long ≤ Q _ Long ( 12 ) where ( Q Lat Q Long) are symmetric positive semi-definite matrices, which can be adjusted as long as uncertainties satisfy the above constraint (12). Note that ξLat and ξLong are unknown time varying matrices satisfying
ξLat T ξLat ≦I and ξLong T ξLong ≦I for and ∀tε[0,∞) (13)
{ U lat = U lat ( X lat ) and V lat = V lat ( X lat ) U long = U long ( X long ) and V long = V long ( X long ) ( 14 ) such that the following norms of transfer functions of longitudinal and lateral constraints are satisfied:
∥G lat(s)+ΔG lat(s)∥≦γlat 2 (15)
∥G long(s)+ΔG long(s)∥≦γlong 2 (16)
G lat = [ ( A lat + A lat ) ( B lat + Δ B lat ) G lat C 1 lat 0 I ( C 2 lat + Δ C 2 lat ) I 0 ] and ( 17 ) G long = [ ( A long + Δ A long ) ( B k long + Δ B long ) G long C 1 long 0 I ( C 2 long ++ Δ C 2 long ) I 0 ] ( 18 ) Note that (γlat, γ long) are small given values, which satisfied constraints (15) and (16). The values (γlat, γlong) are to be determined by the search algorithm, which is described below.
Therefore, to design the required H∞ robust feedback control laws, the H∞ Riccati solutions from the constraints without uncertainties are necessary to be solved first with the good prescribed value selection for (γlat, γlong). Therefore, Riccati solutions, (P∞ lat P∞ long)εdom(H∞), for longitudinal and lateral motions without any uncertainty consideration are associated with the following Hamiltonian matrices,
Note that Qlat=C2 lat T C2 lat and Qlong=C2 long T C2 long, which are symmetric positive semi-definite. The two Riccati inequalities corresponding to the proceeding two Hamiltonian matrices are described as follows:
P ∞ lat A lat +A lat T P ∞ lat +Q lat +P ∞ lat[Ξlat ]P ∞ lat≧0
where Ξlat=γ−2 G lat G lat T −B lat B lat T (21)
P ∞ long A long +A long T P ∞ long +Q long +P ∞ long[Ξlong ]P ∞ long≧0
where Ξlong=γ−2 G long G long T −B long B long T (22)
Note that (Ξlat Ξlong) are symmetric positive definite for prescribed values (γlat γlong). The system satisfies stabilizable and detectable requirements. The solutions of the above Riccati inequalities will be used as initial condition to find the desired solutions for any uncertainties matrices in transfer functions (15) and (16).
U lat =−K lat X lat (23)
U long =−K long X long (24)
The above two Hamiltonian matrices have solutions ( P ∞ lat P ∞ long)εdom(H∞), from the following Riccati inequalities:
P ∞ lat A lat +A lat T P ∞ lat+(Q lat + Q lat)+ P ∞ lat[Ξlat ] P ∞ lat≧0 (27)
P ∞ long A long +A long T P ∞ long+(Q long + Q long)+ P ∞ long[Ξlong ] P ∞ long≧0 (28)
1) (ζlat ζlong): Individual damping ration of closed loop systems
2) (qlat qlong): Diagonal search increment variables for ( Q lat Q long).
With the search algorithm of FIG. 3, the H∞, control gain generated from these steps will make the uncertainty in systems (4) and (5) be stable. Therefore, the H∞ state feedback control gains can be determined as
K lat =−└B lat B lat T ┘ P ∞ lat (29)
K long =−└B long B long T ┘ P ∞ long (29)
1) When a storm is detected, system 27 will command the UAV to reduce or increase its airspeed to maintain the airspeed within a specified safety range.
2) When a storm is detected, system 27 will trigger an automatic altitude recovery function.
The automatic altitude recovery function is dependent on information from (a) an ADC static probes for altitude, (b) a radar altimeter (if installed), (c) a GPS altitude (if still working), (d) transmitted ground station information (if still working), or (e) other sensor information. Usually, microbursts occur at a low altitude, and once the “storm detected” function is triggered, it is necessary for the UAV to fly out of the microburst as soon as possible with the safety airspeed and altitude, total control surface response, and some sensor information.
1) When a storm is detected, system 27 will have an automatic maneuver recovery function.
1) When a storm is detected, system 27 selects ground station intercept mode.
When a storm is detected, system 27 will send out “storm detected” information to a ground station. The ground station can intercept the UAV for any command to bring back the UAV if the satellite signal is still working. However, the UAV will have its own logic to determine the best survival condition from the robust flight control laws.
1) After the storm, the UAV could be damaged and lose all transmitting and receiving capacities.
Under this condition, a “go home” mode will be triggered in which system 27 commands the UAV to fly to a pre-programmed position. The UAV will continue sending its failure signals to the ground station. After escaping from storm, the UAV will stay out of storm for a specified time and then continue to the next mission. System 27 will have an intelligent logic to select its own flight path if ground station signal is not available, the ground station command is not changed, or if the ground station command is not safe.
1) Flight path selection will automatically be triggered if the ground station signal is not available.
The UAV flight path selection will be based on its storm recovery condition. It can intelligently select to continue its own mission, by-pass to next mission, or select “go home” mode. All these are based on sensor information, equipment health, and flight safety.
1. Robust control law development 2. Uncertainty computations from all models 3. Reversed parameter ID estimation for gust covariance estimation 4. Reversed Dryden power spectral density function for gust estimation 5. Computed automatic gust detecting system 6. Intelligent design of UAV to encounter weather condition 7. Intelligent Escape Mode design when gust power density is too strong 8. Lateral alignment design in storm mode 9. Control surface limit hit escape logic design in storm mode 10. Go Home Mode design when sensor failure after the storm attack 11. Minimum fuel destination selection after the storm attack if sensors failed. While described above as being used with an unmanned aircraft, the system of the invention is applicable to all types of aircraft, including manned aircraft. The system of the invention may also incorporate additional features, including override methods for returning control to a pilot.
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No. 11/663,906, issued by the USPTO on Apr. 1, 2009.21Request for Continued Examination dated Feb. 10, 2010.Referenced byCiting PatentFiling datePublication dateApplicantTitleUS8869537 *Aug 27, 2010Oct 28, 2014Rolls-Royce PlcApparatus and method of operating a gas turbine engineUS9108745 *Jun 26, 2014Aug 18, 2015Dassault AviationMethod for detecting a failure of at least one sensor onboard an aircraft implementing an anemo-inertial loop, and associated systemUS9329584 *Dec 15, 2011May 3, 2016International Business Machines CorporationMethod, system and program for constructing a controllerUS20110079015 *Aug 27, 2010Apr 7, 2011Rolls-Royce PlcApparatus and method of operating a gas turbine engineUS20130158695 *Dec 15, 2011Jun 20, 2013International Business MachinesMethod, system and program for constructing a controllerUS20150006019 *Jun 26, 2014Jan 1, 2015Dassault AviationMethod for detecting a failure of at least one sensor onboard an aircraft implementing an anemo-inertial loop, and associated system* Cited by examinerClassifications U.S. Classification701/10, 244/194, 244/189, 244/181, 244/183International ClassificationG05D1/06Cooperative ClassificationG05D1/0623, B64C13/16European ClassificationB64C13/16, G05D1/06B2BLegal EventsDateCodeEventDescriptionJan 11, 2012ASAssignmentOwner name: TEXTRON INNOVATIONS INC., RHODE ISLANDFree format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:BELL HELICOPTER TEXTRON INC.;REEL/FRAME:027515/0353Effective date: 20081211Owner name: BELL HELICOPTER TEXTRON INC., TEXASFree format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:SHUE, SHYHPYNG JACK;REEL/FRAME:027515/0190Effective date: 20050216Nov 2, 2015FPAYFee paymentYear of fee payment: 4RotateOriginal ImageGoogle Home - Sitemap - USPTO Bulk Downloads - Privacy Policy - Terms of Service - About Google Patents - Send FeedbackData provided by IFI CLAIMS Patent Services