Search Images Maps Play YouTube News Gmail Drive More »
Sign in
Screen reader users: click this link for accessible mode. Accessible mode has the same essential features but works better with your reader.


  1. Advanced Patent Search
Publication numberUSH2206 H1
Publication typeGrant
Application numberUS 10/975,112
Publication dateDec 4, 2007
Filing dateOct 28, 2004
Priority dateOct 28, 2004
Publication number10975112, 975112, US H2206 H1, US H2206H1, US-H1-H2206, USH2206 H1, USH2206H1
InventorsJudah H. Milgram
Original AssigneeThe United States Of America As Represented By The Secretary Of The Navy
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Tactile side-slip corrective yaw control for aircraft
US H2206 H1
Side-slip of an aircraft during flight is detected through a pair of pressure sensors fixedly mounted on opposite lateral sides of the aircraft fuselage. Pressure measurement signals at said sensors are fed to electronic circuitry within the aircraft for generating magnitude and frequency signals reflective of the side-slip that are applied to a pair of vibrators respectively mounted on the undersides of a pair of pilot foot pedals located within the cockpit. The foot pedals are connected by linkage to the tail rudder on the aircraft fuselage. The varying magnitude and frequency of vibrations applied to the rudder foot pedals by the vibrators enables the pilot to immediately sense side-slip through the feet on the pedals. In response to such side-slip sensing, one of the pedals may be timely depressed for side-slip corrective angular displacement of the rudder.
Previous page
Next page
1. In combination with an aircraft having a pilot cockpit mounting therein foot-operated means for directional aircraft control about a longitudinal yaw axis by a pilot seated within the cockpit; tactile means for providing pilot perception of side-slip condition, comprising: sensing means for detection of any aircraft side-slip reflected by angular deviation of air flight direction from the longitudinal yaw axis; vibrator means for selectively imparting vibrations directly to the foot-operated means in response to said detection of the side-slip; and signal generating means operatively interconnected between the sensing means and the vibrator means supplying side-slip indicating signals for said perception of the side-slip condition by the pilot through the foot-operated means.
2. The combination as defined in claim 1, wherein said foot-operated means within the cockpit are pilot pedals.
3. The combination as defined in claim 1, wherein said vibrator means comprises: a pair of vibration devices mechanically connected to said foot-operated means.
4. The combination as defined in claim 1, wherein said sensing means comprises: measurement means for measuring alignment of airflow relative to the aircraft longitudinal yaw axis.
5. The combination as defined in claim 4, wherein said measurement means comprises: a pair of static pressure sensing devices fixedly mounted on opposite sides of the aircraft.
6. The combination as defined in claim 1 wherein said sensing means comprises: sensor means fixedly mounted on the aircraft for sensing lateral acceleration thereof.
7. The combination as defined in claim 1 wherein said signal generating means comprises: computational means for computing control inputs based on a mathematical model of aircraft flight mechanics.
8. The combination as defined in claim 1 wherein said signal generating means provides a signal proportional to the side-slip; and modulator means for proportionally modulating the vibrator means to render the vibrations proportional to the side-slip.
9. The combination as defined in claim 1 wherein said sensing means comprises: sensor means for sensing aircraft lateral acceleration; and measurement means for measuring aircraft misalignment of airflow; said signal-generating means being selectably configured either to lateral acceleration of the aircraft or the detection of the side slip.

The invention described herein may be manufactured and used by or for the Government of the United States of America for governmental purposes without the payment of any royalties thereon or therefore.

The present invention relates to aircraft control for side-slip corrective purposes.


Traditionally, aircraft are maneuvered by pitch and yaw control inputs applied in a coordinated manner for smooth directional maneuvering of the aircraft. During such maneuvering of the aircraft, the desired pilot coordinated control over a rudder maintains alignment between the fuselage longitudinal yaw axis and the oncoming flow of air during flight of the aircraft under a zero side-slip condition so as to (a) minimize drag, (b) reduce risk of inadvertent spin under low speed flight, and (c) provide for aircraft passenger comfort. In certain rotary-wing types of aircraft small residual side-slip is desired to counteract lateral load due to the tail rotor by pilot application of a sufficient degree of foot pedal depression. Various automatic maneuvering control systems have however been proposed for establishing the aforesaid desirable coordinated maneuvering control, because of the pilot's inability to continuously provide it manually. Various disadvantages have however been inherently associated with such automatic control systems. It is therefore an important object of the present invention to augment direct pilot maneuvering control by providing immediate tactile perception to the pilot so as to enable corrective response to aircraft side-slip due to non-alignment between the airflow flight path and the yaw axis and thereby avoid any substantial deviation from zero side-slip condition.


Pursuant to the present invention, an aircraft is provided with means for measuring the aircraft side-slip angle. Such side-slip measurements are utilized to generate signals with magnitude and frequency corresponding to side-slip, applied to tactile vibrators mounted on the underside of a pair of pilot foot controls providing the pilot with sense touch perception of any side-slipping condition, thereby enabling immediate pilot depression of the foot controls for corrective control over the aircraft relative to the yaw axis so as to minimize the perceived side-slip. The pilot may thereby provide such corrective control without visual reference to cockpit instruments.


A more complete appreciation of the invention and many of its attendant advantages will be readily appreciated as the same becomes better understood by reference to the following detailed description when considered in connection with the accompanying drawing wherein:

FIG. 1 is a side elevation view of an aircraft under flight, with side-slip corrective maneuvering facilities pursuant to the present invention;

FIG. 2 is a partial fragmentary view of a cockpit portion of the aircraft shown in FIG. 1;

FIG. 3 is a schematic circuit diagram of the tactile side-slip corrective rudder control system associated with the aircraft as shown in FIGS. 1 and 2; and

FIG. 4 is a front elevation view of a helicopter type aircraft with a lateral accelerometer pursuant to another embodiment of the present invention.


Referring now to the drawing in detail, FIG. 1 illustrates a typical aircraft 10 during flight having a longitudinal yaw axis 11. The aircraft 10, which may be of a glider type, has a fuselage 14 with wings 16 attached thereto as well as horizontal and vertical stabilizers 18 and 20 at the tail end thereof. Pivotally connected to the horizontal stabilizers 18 are elevators 22, while a rudder 24 is pivotally connected the vertical stabilizer 20.

Maneuvering of the aircraft 10 as generally known in the art involves displacement of the elevators 22, the rudder 24 and ailerons under control of the pilot in the aircraft fuselage cockpit 26.

As shown in FIG. 2, pilot control is exercised by a pilot when seated on a seat 28 within the cockpit 26 to thereby manually manipulate control yoke 30 and depress a pair of rudder foot control pedals 32 operatively connected by linkages 34 to the rudder 24. Pursuant to the present invention, electromechanical vibrators 36 are mounted on the underside of the petals 32 so as to provide tactile or sense of touch signals applied to the pilot feet 38 positioned on the pedals 32 as depicted in FIG. 3, for yaw axis maneuvering control through the rudder 24 on the tail end of the aircraft fuselage 14. Generation of the pilot foot signals by the vibrators 36 applied to the pilot feet 38 is under control of a system 40 as diagrammed in FIG. 3, which includes a pair of static pressure detectors 42 mounted on opposite port and starboard sides of the fuselage 14.

With continued reference to FIG. 3, the rudder pedal vibrators 36 are connected by electric power cables 46 of the system 40 to a controller 48 which is connected to a side-slip yaw sensor 50 for reception of side-slip air pressure sensing signals received from the side-slip detectors 42. The side-slip detection signals from the detectors 42 is directly fed to a differential transducer 52. The output from the transducer 52 and a transducer 54 is fed to a calibrating control 56 from which supply of calibrated side-slip detection signals are fed to the side-slip sensor 50.

Based on the foregoing description, the controller 48 uses the side-slip signal from the sensor 50 to operate a selected one of the two vibrators 36 so as to signify to the pilot by vibration applied to one of the pilot feet 38 which of the rudder pedals 32 is to be depressed so as to effect angular displacement of the rudder 24 in one direction for side-slip error corrective purposes. The controller 48 may embody a dead band operational feature to by-pass selected pedal vibrator operation when the side-slip error is too small for correction. Furthermore, the controller 48 may incorporate a pilot actuated resetting switch and associated resetting circuit for varying the vibrating pressure applied to the rudder pedals 32 by the vibrators 36 so as accommodate different pilot sensibility preferences. The controller 48 may accordingly be selectively set to generate signals applied to the vibrators 36 by measurements of side-slip to be corrected by precise pilot yaw control through the rudder pedals 32. The signal correction proportionality measurement parameters of the controller 48 may be tuned to side-angle deviation, vibration frequency and vibration magnitude of the vibrators 36. Control over the rudder pedal vibrators 36 through the controller 48 may also be utilized for signifying the requirement of corrective pedal foot input to avoid forthcoming dangerous flight conditions alerted to the pilot.

It will be apparent from the foregoing description that in addition to controlled maneuvering of the aircraft 10 during flight through the steering wheel 30, the pedals 32, the horizontal stabilizer elevators 22 and the rudder 24, as generally known in the art, side-slip corrective adjustment control may be applied to the rudder 24 by the pilot through the pedals 32 in response to tactile sensing of vibrations applied thereto by the vibrators 36. Such vibrations vary in magnitude and frequency in accordance with the detection of aircraft side-slip through the side-slip detector ports 42 and a side-slip indicating vane 44.

FIG. 4 illustrates an asymmetric type of aircraft such as a helicopter 58, which is to be maneuvered during flight with zero lateral acceleration pursuant to another embodiment of the present invention. Laterally mounted on the fuselage 60 of the aircraft 58 is an accelerometer sensor 62 for detecting any lateral side-slip movement of the aircraft 58 which is to be eliminated under a pilot control system involving a control box switchable between zero side-slip mode and ball in the middle mode. Such maneuvering control system associated with the asymmetrical helicopter aircraft 58 may also be applicable to a symmetrical type of aircraft.

Obviously, other modifications and variations of the present invention may be possible in light of the foregoing teachings. It is therefore to be understood that within the scope of the appended claims the invention may be practiced otherwise than as specifically described.

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2008693 *Nov 28, 1931Jul 23, 1935Fator Charles DSignaling system
US2442289Apr 6, 1945May 25, 1948Edward YoungAirplane control system
US2697566Oct 11, 1949Dec 21, 1954Boeing CoSelective two or three control type system for aircraft
US3076624Nov 16, 1960Feb 5, 1963U S Science CorpOscillatory alarm for aircraft and the like
US3792426 *Apr 6, 1972Feb 12, 1974Us Air ForceTactile warning device for g-loading angle of attack
US3902687 *Jun 25, 1973Sep 2, 1975Robert E HightowerAircraft indicator system
US4195802Apr 28, 1978Apr 1, 1980The Ohio State UniversityKinesthetic tactile display system
US4206891Oct 26, 1978Jun 10, 1980United Technologies CorporationHelicopter pedal feel force proportional to side slip
US4484191 *Jun 14, 1982Nov 20, 1984Vavra George STactile signaling systems for aircraft
US4814764 *Sep 30, 1986Mar 21, 1989The Boeing CompanyApparatus and method for warning of a high yaw condition in an aircraft
US5062594Nov 29, 1990Nov 5, 1991The United States Of America As Represented By The Secretary Of The Air ForceFlight control system with tactile feedback
US5467322Oct 26, 1994Nov 14, 1995Ind Sound Technologies IncWater hammer driven vibrator
US5738310Dec 22, 1995Apr 14, 1998Eurocopter FranceRudder bar system with force gradient for a helicopter
US5852237 *May 28, 1997Dec 22, 1998Lockheed Martin CorporationApparatus and method for measuring the side slip of a low observable aircraft
US6002349 *Aug 14, 1998Dec 14, 1999Safe Flight Instrument CorporationHelicopter anti-torque limit warning device
US6253126 *May 12, 1998Jun 26, 2001Aers/Midwest, Inc.Method and apparatus for flight parameter monitoring and control
US6273371 *Nov 10, 1999Aug 14, 2001Marco TestiMethod for interfacing a pilot with the aerodynamic state of the surfaces of an aircraft and body interface to carry out this method
Non-Patent Citations
1Craig, G., Jennings, S., Cheung, B, Rupert, A., Schultz, K., "Flight-Test fo a Tactile Situational Awardness System in a high-Hover Task", American Helicopter Soc. 60th Annual Forum, Baltimore, Maryland, June 7-10, 2004, 7 pages.
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US8584990 *Feb 9, 2012Nov 19, 2013Airbus Operations (Sas)Method and device for yaw controlling of an aircraft
US8718841Feb 14, 2012May 6, 2014Sikorsky Aircraft CorporationMethod and system for providing sideslip envelope protection
US20120205495 *Feb 9, 2012Aug 16, 2012Airbus Operations (S.A.S.)Method And Device For Yaw Controlling Of An Aircraft
U.S. Classification244/75.1, 244/76.00R, 244/194, 244/196, 244/195
International ClassificationB64C13/24
Cooperative ClassificationB64C13/46
European ClassificationB64C13/46
Legal Events
Nov 15, 2004ASAssignment
Effective date: 20041025