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Publication numberUS2532483 A
Publication typeGrant
Publication date5 Dec 1950
Filing date2 Jan 1948
Priority date2 Jan 1948
Publication numberUS 2532483 A, US 2532483A, US-A-2532483, US2532483 A, US2532483A
InventorsElliot Daland, Meyers Donald N
Original AssigneePiasecki Helicopter Corp
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Control for tandem rotored helicopters
US 2532483 A
Abstract  available in
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Claims  available in
Description  (OCR text may contain errors)

E. DALAND EIAL 2,532,483

coumor. FOR TANDEM ROTORED HELICOPTERS Dec. 5, 19 50 2 Sheets-Sheet 1 Filed Jan. 2, 1948 Dec. 5, 1950 E. DALAND ETAL CONTROL FOR TANDEM ROTORED HELICOPTERS 2 Sheets-Sheet 2 Filed Jan. 2, 1948 INVEN T RS M T @a bn-d,

Patented Dec. 5, "1950 CONTROL FOR TANDEM ROTORED HELICOPTERS Elliot Daland, Walllngford, and Donald N. Meyers, Philadelphia, Pa., assignors, by mesne-assignments, to

Morton, Pa., a corporation Piasecki Helicopter Corporation,

of Pennsylvania Application January 2, 1948, Serial No. 216

3 Claims.

the pilot decreases the collective pitch of thefront rotor and increases the collective pitch of the rear rotor simultaneously, by moving"- the control column, provided for this purpose, forwardly. To tilt the machine upward, th'control column is moved rearwardly by the pilot, simultaneously increasing the collective pitch of the front rotor and decreasing the collective pitch of the rear rotor.

Directional control is. obtained by oppositely tilting the fore and aft rotors sideways by the use of cyclic pitch. Rudder pedals are provided for pilot operation to difierentially apply cycliccontrol to the rotors to turn the machine to the right or left.

The rotors on this type of machine are turned in opposite directions to neutralize the torque reaction acting on the craft. We have found from actual practice that turning the rotors in opposite directions does not entirely neutralize torque reaction forces and a rudder correction must be made to maintain a straight course. This results in a partial loss of rudder control in the direction that the rudder correction is being appii i. We have also learned from practic that when helicopters of this type are in forward flight attitude and the rotor blades of both rotors a-e set at the same incidence angle as has been common practice, the control column must be carried in a position ahead of mid-position :to' simultaneously increase the collective pitch-'of the rear rotor and decrease the collective pitch of the front rotor in order to equalize the lift of the two rotors. This limits the am unt of forward control available and requires the use of additional total collective pitch.

The rotor blades of helicopters provided with the aforedescribed control systems have been rigged in such a manner that the blades of both of the rotors are set at the same incidence angle when the control is in mid-position; and the directional control system has been rigged so that when the swash plates are horizontal, the rudder pedals are in a neutral position.

Accordingly, it is among the principal objects of the present inventicn to provide a method of rigging the rotor blades in the control system to (Cl. ran-135.24)

. 2 eliminate the climbing moment acting on the helicopter when in forward flight attitude and offset the turning forces necessitating the use of therudder control.

Another object of this invention is the equalization of the degree of control that may be exercised by the use of the rudder pedals or other directional means provided.

Another object of this invention is to equalize the thrust of the two rotors by setting the blades of the rear rotor at a higher incidence angle than those of the front rotor when the control column is in a mid-position.

A further object of this invention is to provide a greater degree of control about the yawing and pitching ax'es of the aircraft.

The method of rigging as provided by our invention overcomes the faults and disadvantages attributable to former methods of rigging and obtains the objects and advantages outlined above.

White this invention may be applied to all heli copters of the tandem type, we have chosen for purposes of illustration, to show its application to the control system disclosed and described in the copending application by Elliot Daland Serial No. 729,565, filed February 19, 1947, the drawings of which are reproduced herein with slight modification to better illustrate this invention.

In the drawings, Figure 1 is an isometric view of the forward portion of the controls and a partial view of the front rotor.

Figure 2 is an isometric view of the aft portion of the controls and a partial view of the rear rotor.

A full description of all the details shown in the drawings willnot be given here as they are described in the afore-mentioned pending applicaion.

Referring more particularly to the drawings, the control system is comprised of a lcngitudinal and lateral control column II, a total collective pitch control 12, and directional control pedals l3 and Il. These pilot actuated controls are mechanically connected, to the swash plates l5 and H; which act to change the pitch of the rotor blades l1 and I8. It is understood of course that two or more blades of the type shown are usually employed in each rotor.

Forward movement of the control I I will simultaneously increase the pitch of the rear rotor blades l8 and decrease the pitch of the front rotor blades l1. Moving the control column II to the right will tilt the swash plates l5 and It to the right, cyclically changing the pitch of both rotors to bank the ship to the right. To bank the ship to the left the control II is moved to the left. Backward movement of the control column ll increases the collective pitch of the front rotor and decreases the collective pitch of the rear rotor. Pushing the right rudder pedal l3 tilts accuse the front swash plate I and lift vector of the forward rotor to the right and tilts the aft swash plate I 6 and rear rotor lift vector to the left tuming the machine to the right. Pushing the left rudder pedal I4 causes an opposite effect. Upward,movement of the control l2 increases the pitch of both rotors collectively. Downward 'movement of the control I 2 decreases the pitch of both rotors collectively.

Ourimproved method of rig ing consists of setting the front rotor blades I! at a predetermined incidence angle when the longitudinal and total collective pitch controls II and I2 are in their mid-positions.

As the rear motor is acting in the downwash of the front rotor when the helicopter is in forward flight attitude or hovering into a wind, the angle of attack of the rear rotor blades 18 will be less than that of the front rotor blades l1 resulting in unequal lift distribution between the two rotors. To overcome this condition we set the blades 18 of the rear rotor at a higher incidence angle than those of the front rotor. This is 'done with the longitudinal and total collective pitch controls set in their mid-position.

The term incidence angle as used herein defines the relative angle between the chord line of the blade and the plane of rotationof the rotor. This adjustment is madewithout disturbing the vertical settings of the swash plates is and I6 by rotating the blades about their pitch axes. By doing this the thrust of the two rotors is equalized when the craft is flying at approximately its cruising speed with the total collective pitch and longitudinal control columns in their mid-positions. 2

It is now apparent from the foregoing description that this method of rigging provides the connecting the pilot actuated means to said blade pitch changing means to simultaneously increase the pitch of the blades of one rotor and decrease the pitch of blades of the other rotor on imparting one kind of movement to the pilot actuated means, said connecting means and rotor blades being so adjusted that the angle of incidence of the rear rotor blades is greater than the angle of incidence of the forward rotor blades when the pilot actuated means is at the center of travel of said one kind of movement.

2. A helicopter control system comprising two equivalent, counter-rotating lift rotors mounted in tandem, each of said rotors havingpne or more blades and being provided with blade pitch changing means, pilot actuated means movable between two limiting positions for controlling the helicopter about its yaw axis, means connecting the pilot actuating means to said blade pitch changing means to oppositely tilt the thrust vectors of said rotors laterally on movement of said pilot actuated means, said connecting means and blade pitch changing means being so adjusted that the lift vectors of the rotors are oppositely tilted in a lateral direction when the pilot actuated means is centered with respect to said limiting positions.

3. A helicopter control system comprising two equivalent lift rotors mounted in tandem, blade pitch changing means located adjacent to each of said rotors, the blades of said rotors being connected to said blade pitch changing means for pitch adjustment, a first pilot control means, means connecting said pilot control means to said blade pitch changing means to simultaneously increase the pitch of one rotor and decrease the greatest degree of control each side of the mid positions of the control members thereby eliminating the danger of running out of control as has happened with other systems of control rig- Because of the aforementioned downwash effect 0n the rear rotor and the geometry of the lift forces about the center of gravity of the machine, we have found that the rear rotor requires more power than that of the front rotor which results in a turning moment being set up about the yawing axis of the craft caused by the unequal torque reaction forces acting on the craft. To offset this turning moment we have rigged the directional control system in such a manner that when the rudder pedals I 3 and ll are in a neutral position the two swash plates I5 and 16 are tilted in the opposite directions as shown to apply directional cyclic control of the rotors to overcome this turning moment. This setting is obtained by adjusting the turnbuckles l9 and 20 in such a manner that the left rudder pushpull tube is lengthened and the right rudder push-pull tube is shortened.

Rigging the directional control in this manner eliminates the use of the rudder correction when the controls of the aircraft are in a cruising position and gains the advantage of allowing a, full amount of rudder travel for control purposes.

We claim:

1. A helicopter control system comprising two equivalent lift rotors mounted in tandem, each of said rotors having one or more blades and being provided with blade pitch changing mea pilot actuated means for controlling the attitude of the helicopter about its pitching axis, means pitch of the other rotor, on imparting one kind of movement to said pilot control means, a second pilot controlmeans movable between two limiting positions, and means connecting said second control means and said blade pitch changin means to oppositely tilt the thrust vectors of said rotors laterally upon actuation of said sec- -ond control means by the pilot, the connecting 'means between the flrst mentioned pilot control means and said blade pitch changing means being so proportioned and adjusted that the blades of the rear rotor are at a higher incidence angle than those of the front rotor when the'first-mentioned pilot control means is at the center of travel of said one kind of movement, the connecting means between the second pilot control means and the blade pitch changing means being so proportioned and adjusted that the rotor thrust vectors are oppositely tilted laterally when the second-mentioned control means is centered with respect to said limiting positions, whereby to counteract the unbalanced torque of the rotors.

, ELLIOT DALAND'.

DONALD N. MEYEBS.

REFERENCES CITED The following references are of record in the file of this patent:

UNITED STATES PATENTS 155,914 Switzerland Oct. 1, 1932

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2373575 *3 Oct 194210 Apr 1945Reconstruction Finance CorpControl mechanism for aircraft
CH155974A * Title not available
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US2753004 *15 Mar 19523 Jul 1956Mcculloch Motors CorpPitch control means for aircraft sustaining rotors
Classifications
U.S. Classification416/115, 416/130
International ClassificationB64C27/54
Cooperative ClassificationB64C27/54, B64C2700/6284
European ClassificationB64C27/54