GB2090214A - Controlling Helicopter Rotors - Google Patents

Controlling Helicopter Rotors Download PDF

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Publication number
GB2090214A
GB2090214A GB8026309A GB8026309A GB2090214A GB 2090214 A GB2090214 A GB 2090214A GB 8026309 A GB8026309 A GB 8026309A GB 8026309 A GB8026309 A GB 8026309A GB 2090214 A GB2090214 A GB 2090214A
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GB
United Kingdom
Prior art keywords
rotor
blades
hub
controlled
control
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
GB8026309A
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GB2090214B (en
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Individual
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Individual
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
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Priority to GB8026309A priority Critical patent/GB2090214B/en
Publication of GB2090214A publication Critical patent/GB2090214A/en
Application granted granted Critical
Publication of GB2090214B publication Critical patent/GB2090214B/en
Expired legal-status Critical Current

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C27/00—Rotorcraft; Rotors peculiar thereto
    • B64C27/54—Mechanisms for controlling blade adjustment or movement relative to rotor head, e.g. lag-lead movement
    • B64C27/72—Means acting on blades
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C27/00—Rotorcraft; Rotors peculiar thereto
    • B64C27/54—Mechanisms for controlling blade adjustment or movement relative to rotor head, e.g. lag-lead movement
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B64—AIRCRAFT; AVIATION; COSMONAUTICS
    • B64C—AEROPLANES; HELICOPTERS
    • B64C27/00—Rotorcraft; Rotors peculiar thereto
    • B64C27/54—Mechanisms for controlling blade adjustment or movement relative to rotor head, e.g. lag-lead movement
    • B64C27/58—Transmitting means, e.g. interrelated with initiating means or means acting on blades
    • B64C27/59—Transmitting means, e.g. interrelated with initiating means or means acting on blades mechanical
    • B64C27/615—Transmitting means, e.g. interrelated with initiating means or means acting on blades mechanical including flaps mounted on blades
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/30—Wing lift efficiency

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Toys (AREA)

Abstract

Helicopter rotor hub 10 with the blades as an assembly, the blades can be integral or separate, and controlled without any mechanical control linkage attached to the airframe. Electrical signals and/or optical signals passed through a transmitter unit 16 to the rotor hub position the rotor blades by driving a positioning motor 18 attached to each blade, the motor acting on the blade root, via gearing, or on an aileron on the blade. Electrical power fed to rotors via slip ring 8. An alternative semi- hydraulic control system is described (Fig. 7 not shown). <IMAGE>

Description

SPECIFICATION Laser Electronically Controlled Helicopter Main Rotor Hub and Aileron Controlled Rotor Blades (Rotating Wing Aircraft) My invention will almost completely eliminate rotor induced vibration from helicopters by removing the present accepted method of controlling rotor blades through pitch change control rods to the blades.
Existing systems of changing/controlling pitch all rely on a mechanical linkage, this means that the force required to twist a rotor bide must use the chassis or airframe as the basis for the initial thrust to the rotor blades. This must and in fact does, create a mechanical reaction between airframe (helicopter) and rotor blade directly proportional to the force applied combined with the setting of the blades. For example if the blades are set in collective pitch (hovering flight) vibration will be less for any given revolution per minute, than for the equivalent R.P.M. in cyclic pitch (moving flight).
To take these facts a stage further, it is known fact that more rotor blades equal more vibration.
The proof of this can be seen by the success of existing 2 rotor designs, which are both cheaper to make and maintain and produce less vibration as well.
With my rotor head there will be less vibration than the 2 blade type whilst using any combination of blades a particular design would require. Less complexity will help reduce maintenance times and because of reduced vibration, airframe life will increase as will servicing intervals. Further advantages will acrue from less complexity because of an overall reduction in weight, this will improve payload/range and speed performance figures.
Another benefit, though less important will come from less aerodynamic drag to a much cleaner rotor hub design externally and mechanically because the rotor power unit(s) would only have to rotate the hub and not overcome the mechanical drag associated with conventional pitch change mechanisms.
Another invention is shown incorporated on the drawings enclosed (sheet 3) that is the control of the rotor blade(s) by aileron, exploiting the phenomena known as aileron reversal (on fixed wing aircraft) with the rotor blade rigidly connected to the hub (rigid rotor) or by electrical drive from the hub see sheet 4.
Although partly mechanical the last system will still achieve the sought after reduced vibration levels because the source of pitch change is connected to the rotor hub and will therefore riot transmit mechanically induced vibration through the airframe.
* The use of the LASER in this invention does not preclude the use of alternative electronics to achieve the same control functions and as such must not invalidate my rights to invention and PATENT.
Reference numbers in the drawings (Figs. 19).
1 Blade control cilerons 3 Control unit in cockpit (bias unit incorporated to achieve full cyclic/collective control electronically) 6 Signal wires and power carrying lines 7 Control pulses/signals to aileron position motor 8 Slip rings (in top of rotor hub to transfer electrical power) 9 Control unit (triggerhead) 10 Rotor hub (forged as one piece unit; or moulded/laminated in plastic, e.g. GRP, carbon fibre; blades could be made detachable.
11 Position motor 1 2 Receiving matrix 13 Optional servo tab operation for aileron 14 Receiver 1 5 Laser tubes (fibre optics) 1 6 Matrix of laser transmitter leads 1 8 Laser controlled electric pitch change motor 1 9 Rotor hub (modified conventional rigid hub) 20 Gear driven blade root/s 21 Mechanically controlled transmitter ring 22 Hydraulic positioning arms controlled by unit in cockpit.
Claims
1. An electronically controlled rotor head which will dispense with all direct control linkages to the airframe. See Figs. 1,2 and 3.
2. Aileron controlled rotor blades using rotor twist or flexion to achieve pitch change/control.
See Fig. 1.
3. Rotor blade pitch controlled by electric servo motor(s) controlled from the triggerhead tube.
See Fig. 2.
**WARNING** end of DESC field may overlap start of CLMS **.

Claims (1)

  1. **WARNING** start of CLMS field may overlap end of DESC **.
    SPECIFICATION Laser Electronically Controlled Helicopter Main Rotor Hub and Aileron Controlled Rotor Blades (Rotating Wing Aircraft) My invention will almost completely eliminate rotor induced vibration from helicopters by removing the present accepted method of controlling rotor blades through pitch change control rods to the blades.
    Existing systems of changing/controlling pitch all rely on a mechanical linkage, this means that the force required to twist a rotor bide must use the chassis or airframe as the basis for the initial thrust to the rotor blades. This must and in fact does, create a mechanical reaction between airframe (helicopter) and rotor blade directly proportional to the force applied combined with the setting of the blades. For example if the blades are set in collective pitch (hovering flight) vibration will be less for any given revolution per minute, than for the equivalent R.P.M. in cyclic pitch (moving flight).
    To take these facts a stage further, it is known fact that more rotor blades equal more vibration.
    The proof of this can be seen by the success of existing 2 rotor designs, which are both cheaper to make and maintain and produce less vibration as well.
    With my rotor head there will be less vibration than the 2 blade type whilst using any combination of blades a particular design would require. Less complexity will help reduce maintenance times and because of reduced vibration, airframe life will increase as will servicing intervals. Further advantages will acrue from less complexity because of an overall reduction in weight, this will improve payload/range and speed performance figures.
    Another benefit, though less important will come from less aerodynamic drag to a much cleaner rotor hub design externally and mechanically because the rotor power unit(s) would only have to rotate the hub and not overcome the mechanical drag associated with conventional pitch change mechanisms.
    Another invention is shown incorporated on the drawings enclosed (sheet 3) that is the control of the rotor blade(s) by aileron, exploiting the phenomena known as aileron reversal (on fixed wing aircraft) with the rotor blade rigidly connected to the hub (rigid rotor) or by electrical drive from the hub see sheet 4.
    Although partly mechanical the last system will still achieve the sought after reduced vibration levels because the source of pitch change is connected to the rotor hub and will therefore riot transmit mechanically induced vibration through the airframe.
    * The use of the LASER in this invention does not preclude the use of alternative electronics to achieve the same control functions and as such must not invalidate my rights to invention and PATENT.
    Reference numbers in the drawings (Figs. 19).
    1 Blade control cilerons
    3 Control unit in cockpit (bias unit incorporated to achieve full cyclic/collective control electronically)
    6 Signal wires and power carrying lines
    7 Control pulses/signals to aileron position motor
    8 Slip rings (in top of rotor hub to transfer electrical power)
    9 Control unit (triggerhead)
    10 Rotor hub (forged as one piece unit; or moulded/laminated in plastic, e.g. GRP, carbon fibre; blades could be made detachable.
    11 Position motor 1 2 Receiving matrix
    13 Optional servo tab operation for aileron
    14 Receiver 1 5 Laser tubes (fibre optics) 1 6 Matrix of laser transmitter leads 1 8 Laser controlled electric pitch change motor 1 9 Rotor hub (modified conventional rigid hub)
    20 Gear driven blade root/s
    21 Mechanically controlled transmitter ring
    22 Hydraulic positioning arms controlled by unit in cockpit.
    Claims
    1. An electronically controlled rotor head which will dispense with all direct control linkages to the airframe. See Figs. 1,2 and 3.
    2. Aileron controlled rotor blades using rotor twist or flexion to achieve pitch change/control.
    See Fig. 1.
    3. Rotor blade pitch controlled by electric servo motor(s) controlled from the triggerhead tube.
    See Fig. 2.
GB8026309A 1980-08-13 1980-08-13 Controlling helicopter rotors Expired GB2090214B (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
GB8026309A GB2090214B (en) 1980-08-13 1980-08-13 Controlling helicopter rotors

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB8026309A GB2090214B (en) 1980-08-13 1980-08-13 Controlling helicopter rotors

Publications (2)

Publication Number Publication Date
GB2090214A true GB2090214A (en) 1982-07-07
GB2090214B GB2090214B (en) 1984-09-12

Family

ID=10515419

Family Applications (1)

Application Number Title Priority Date Filing Date
GB8026309A Expired GB2090214B (en) 1980-08-13 1980-08-13 Controlling helicopter rotors

Country Status (1)

Country Link
GB (1) GB2090214B (en)

Cited By (17)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4655685A (en) * 1985-12-16 1987-04-07 United Technologies Corporation Helicopter main rotor blade having a short span slot near the tip end
DE3614371A1 (en) * 1986-04-28 1987-10-29 Henschel Flugzeug Werke Gmbh Device for controlling rotor blades
US4720059A (en) * 1986-12-31 1988-01-19 Stearns Jr Hoyt A High speed helicopter
US4808072A (en) * 1987-09-17 1989-02-28 United Technologies Corporation Electrical grounding connector for a helicoptor rotor
US4946354A (en) * 1987-12-15 1990-08-07 Aerospatiale Societe Nationale Industrielle Hydraulic device for individual control of pitch of a rotor blade
GB2280412A (en) * 1993-07-15 1995-02-01 Nigel Howard Mckrill Helicopter rotor controlled by light signals.
WO1996001503A1 (en) * 1994-07-06 1996-01-18 Mcdonnell Douglas Helicopter Company Piezoelectric actuator
GB2298624A (en) * 1995-03-04 1996-09-11 Nigel Howard Mckrill Light controlled helicopter rotor blade
GB2299562A (en) * 1995-04-01 1996-10-09 Nigel Howard Mckrill Actuator for helicopter rotor blade aileron
WO2005100154A1 (en) * 2004-04-13 2005-10-27 Wavefront Technology Pty Ltd System for rotor head and rotor blade
EP1630097A1 (en) * 2004-08-31 2006-03-01 The Boeing Company Rotor blade with structurally-integrated conduit assembly and method for its use
DE102005007129A1 (en) * 2005-02-17 2006-08-24 Zf Friedrichshafen Ag Arrangement for controlling of rotor blades of helicopter has controlled actuator, for collective and cyclic adjustment of blade setting angle of rotor blades, which are mechanically forcibly coupled among themselves
CN102811905A (en) * 2010-03-23 2012-12-05 贝尔直升机泰克斯特龙公司 Swashplate-mounted permanent magnet alternator
EP2829471A1 (en) * 2013-07-23 2015-01-28 Sikorsky Aircraft Corporation Swashplateless coaxial rotary wing aircraft
DE102015225098A1 (en) * 2015-12-14 2017-06-14 Zf Friedrichshafen Ag Rotor head cover with integrated electronics for a helicopter
CN113928540A (en) * 2021-11-19 2022-01-14 中国直升机设计研究所 Helicopter inertia variable-pitch rotor wing
CN116424552A (en) * 2023-05-26 2023-07-14 北京航空航天大学 Active torsion blade vibration control method, system, equipment and medium

Cited By (26)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4655685A (en) * 1985-12-16 1987-04-07 United Technologies Corporation Helicopter main rotor blade having a short span slot near the tip end
DE3614371A1 (en) * 1986-04-28 1987-10-29 Henschel Flugzeug Werke Gmbh Device for controlling rotor blades
US4720059A (en) * 1986-12-31 1988-01-19 Stearns Jr Hoyt A High speed helicopter
US4808072A (en) * 1987-09-17 1989-02-28 United Technologies Corporation Electrical grounding connector for a helicoptor rotor
US4946354A (en) * 1987-12-15 1990-08-07 Aerospatiale Societe Nationale Industrielle Hydraulic device for individual control of pitch of a rotor blade
GB2280412A (en) * 1993-07-15 1995-02-01 Nigel Howard Mckrill Helicopter rotor controlled by light signals.
WO1996001503A1 (en) * 1994-07-06 1996-01-18 Mcdonnell Douglas Helicopter Company Piezoelectric actuator
US5626312A (en) * 1994-07-06 1997-05-06 Mcdonnell Douglas Corporation Piezoelectric actuator
GB2298624A (en) * 1995-03-04 1996-09-11 Nigel Howard Mckrill Light controlled helicopter rotor blade
GB2299562A (en) * 1995-04-01 1996-10-09 Nigel Howard Mckrill Actuator for helicopter rotor blade aileron
WO2005100154A1 (en) * 2004-04-13 2005-10-27 Wavefront Technology Pty Ltd System for rotor head and rotor blade
JP2006069535A (en) * 2004-08-31 2006-03-16 Boeing Co:The Longitudinal rotor blade, aircraft, and method for operating rotor-driven aircraft
EP1630097A1 (en) * 2004-08-31 2006-03-01 The Boeing Company Rotor blade with structurally-integrated conduit assembly and method for its use
DE102005007129A1 (en) * 2005-02-17 2006-08-24 Zf Friedrichshafen Ag Arrangement for controlling of rotor blades of helicopter has controlled actuator, for collective and cyclic adjustment of blade setting angle of rotor blades, which are mechanically forcibly coupled among themselves
DE102005007129B4 (en) * 2005-02-17 2008-07-10 Zf Friedrichshafen Ag Device for controlling rotor blades of a helicopter
CN102811905B (en) * 2010-03-23 2015-12-02 贝尔直升机泰克斯特龙公司 A kind of tilting frame system and adopt the method for this tilting frame system generation current
CN102811905A (en) * 2010-03-23 2012-12-05 贝尔直升机泰克斯特龙公司 Swashplate-mounted permanent magnet alternator
US9248909B2 (en) 2013-07-23 2016-02-02 Sikorsky Aircraft Corporation Swashplateless coaxial rotary wing aircraft
EP2829471A1 (en) * 2013-07-23 2015-01-28 Sikorsky Aircraft Corporation Swashplateless coaxial rotary wing aircraft
DE102015225098A1 (en) * 2015-12-14 2017-06-14 Zf Friedrichshafen Ag Rotor head cover with integrated electronics for a helicopter
US10351228B2 (en) 2015-12-14 2019-07-16 Zf Friedrichshafen Ag Rotor head with integrated electronics for a helicopter
DE102015225098B4 (en) 2015-12-14 2021-08-05 Zf Friedrichshafen Ag Rotor head cover with integrated electronics for a helicopter
CN113928540A (en) * 2021-11-19 2022-01-14 中国直升机设计研究所 Helicopter inertia variable-pitch rotor wing
CN113928540B (en) * 2021-11-19 2023-10-27 中国直升机设计研究所 Helicopter inertia displacement rotor wing
CN116424552A (en) * 2023-05-26 2023-07-14 北京航空航天大学 Active torsion blade vibration control method, system, equipment and medium
CN116424552B (en) * 2023-05-26 2023-08-11 北京航空航天大学 Active torsion blade vibration control method, system, equipment and medium

Also Published As

Publication number Publication date
GB2090214B (en) 1984-09-12

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Legal Events

Date Code Title Description
PCNP Patent ceased through non-payment of renewal fee

Effective date: 19970813