Aeronautics and astronautics – Aircraft sustentation – Sustaining airfoils
Reexamination Certificate
2003-01-03
2004-02-03
Barefoot, Galen (Department: 3644)
Aeronautics and astronautics
Aircraft sustentation
Sustaining airfoils
C199S075000, C199S028000
Reexamination Certificate
active
06685143
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates to a forebody flow control system and more particularly to aircraft or missile flow control system for enhanced maneuverability and stabilization at high angles of attack. The present invention further relates to a method of operating the flow control system.
2. Technical Background
In numerous aeronautical applications it is desirable to control the flow across a surface. As fluid flows over a flow surface, like air over an aircraft or a missile fore body, it forms a fluid boundary layer at the surface. The fluid boundary layer is a thin layer of viscous flow exhibiting certain pressure variations that affect the operation of the aircraft or a missile.
One of these variations is the separation and vortex induced phantom yaw caused by asymmetric vortex shedding on an aircraft or a missile at high angles of attack, even at zero angle of sideslip of. Large forces and dynamic out-of-plane loading on the aircraft or missile occur at angles of attack ranging from 30 to 60 degrees. It is known that the out-of-plane loading results from micro-asymmetries on the surface of the nose of the aircraft or missile such as dents, cracks in the paint and other microscopic imperfections near the tip of the nose. It has also been known that these asymmetries are affected by the bluntness of the forebody, Reynolds Number; roll angle, and the angle of attack. At high angles of attack, these side forces (yaw) are especially pronounced due to ineffectiveness of the traditional flight control surfaces. Side forces resulting from these asymmetries adversely affect the missile or aircraft's performance and significantly limit their flight envelope.
The demand for better control of missiles or aircraft at high angles of attack has led to a number of approaches for control of these side forces. Flow control devices have been employed to control and counteract these side forces. These flow control devices are either passive or active. Passive flow control devices have included geometric changes to the forebody structure such as nose bluntness, strakes, boundary layer strips, vane vortex generators and rotating nose tips to control the asymmetric vortices off the forebody. These passive flow control techniques are effective to some extent in alleviating these side forces, but at the same time limit the performance of the aircraft or missile by increasing the drag. Active flow control devices have included jet blowing, unsteady bleed, suction, blowing and deployable flow effectors to control the asymmetric vortices off the fore body. These active flow control techniques are (as with passive devices) also effective to some extent in alleviating these side forces, but also not optimized (as with passive devices), because they operate in an open-loop mode with no sensor feedback, at the same time limit the performance of the aircraft or missile by increasing the drag.
In view of the foregoing disadvantages with presently available passive or active flow control systems and methods for controlling flow asymmetries on a missile or an aircraft, it has become desirable to develop a missile or aircraft forebody flow control system that controls both the magnitude and direction of these side forces (and further the aircraft or missile maneuverability), and can be deactivated when not required in order to reduce drag.
SUMMARY OF THE INVENTION
The present invention relates to a forebody flow control system and more particularly to aircraft or missile flow control system for enhanced maneuverability and stabilization at high angles of attack. The present invention further relates to a method of operating the flow control system.
In one embodiment, the present invention includes a missile or aircraft comprising an afterbody and a forebody; at least one deployable flow effector on the missile or aircraft forebody; at least one sensors each having a signal, the at least one sensor being positioned to detect flow separation on the missile or aircraft forebody; and a closed loop control system; wherein the closed loop control system is used for activating and deactivating the at least one deployable flow effector based on at least in part the signal of the at least one sensor.
In another embodiment, the present invention includes a flow control system for a missile or aircraft forebody comprising at least one activatable flow effectors; at least one sensor having a signal, the at least one sensor being positioned to detect flow separation on the missile or aircraft forebody; an inertial measurement unit having an output; and a closed loop control system; wherein the closed loop control system is used for activating and deactivating the at least one flow effector based on at least in part the signal of the at least one sensor and the output of the inertial measurement unit.
In still another embodiment, the present invention includes a method of stabilization for a missile or aircraft forebody comprising the steps of estimating or determining side forces on a missile or an aircraft forebody based at least in part on a signal from at least one sensor, the at least one sensor being positioned to detect flow separation on the missile or aircraft forebody; the missile or aircraft forebody further comprising at least one flow effector and a closed loop control system for controlling the flow effectors; activating the at least one flow effectors to counteract the side forces by oscillation of the at least one flow effector with the closed loop controller based on at least in part the signal of the at least sensor; and re-estimating or determining side forces on the missile or aircraft forebody based at least in part on a signal from the at least one sensor; and deactivating the at least one flow effector in response to reduced or changed side forces.
Additional features and advantages of the invention will be set forth in the detailed description which follows, and in part will be readily apparent to those skilled in the art from that description or recognized by practicing the invention as to described herein, including the detailed description which follows, the claims, as well as the appended drawings.
REFERENCES:
patent: 4917333 (1990-04-01), Murri
patent: 5755408 (1998-05-01), Schmidt et al.
patent: 6105904 (2000-08-01), Lisy et al.
Lars Ericsson and Martin Beyers; Forebody Flow Control at Conditions of Naturally Occuring Separation Asymmetry; Journal of Aircraft, Vol. 39, No. 2, Mar.-Apr. 2002; pp 252-261.
L.E. Ericsson and J.P. Reding; Asymmetric Flow Separation and Vortex Shedding on Bodies of Revolution; From: Tactical Missile Aerodynamics: General Topics Edited by Michael J. Hemsch; Vol 141, Chapter No. 10; 1989; pp 391-401.
Lisa Bjarke, John Frate, and David Fisher; A Summary of the Forebody High-Angle-of-Attack Aerodynamics Research on the F-18 and the X-29A Aircraft; Nasa Technical Memorandum, Nov. 1992; pp 1-17.
David Fisher and Daniel Murri; Forebody Flow Visualization on the F-18 HARV With Actuated Forebody Strakes; Nasa Technical Memorandum, Sep. 1998; pp 1-10.
Mehul Patel, Troy Prince, Reed Carver, Jack DiCocco, Frederick Lisy, and Terry Ng; Deployable Flow Effectors for Phantom Yaw Control of Missiles at High Alpha; 1st AIAA flow Control Conference Jun. 24-26, 2002, St Louis, MO; pp 1-12.
J.E. Bernhardt and D.R. Williams; Closed Loop Control of Forebody Flow Asymmetry; Journal of Aircraft vol. 27, No. 3; May-Jun. 2000; pp 491-498.
Carver Reed
DiCocco Jack M.
Lisy Frederick J.
Patel Mehul P.
Prince Troy
Barefoot Galen
Kolkowski Brian M.
Orbital Research Inc.
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