JPS5951479B2 - Control method and aircraft for aircraft with twin-rotating hingeless rotor wings - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION The present invention relates to the flight control of helicopters, and more particularly to the flight control of helicopters having coaxial twin-rotating rotors for optimum performance and maneuverability.

Although it has been known for a long time that advantages can be obtained by adopting a coaxial twin-rotating type hingeless rotor blade as a lift generation mechanism for a helicopter, there were several problems including control problems. The development of actual helicopters employing rotary wing systems such as those described above has only recently occurred.

Igor I, a famous aviation pioneer.

It is noteworthy that when Ikorsky built his first helicopter at the beginning of the 20th century, it included a coaxial twin-rotating hingeless rotor.

Glauert wrote his book “Aerodynamic
In ``Theory,'' he suggested that the roll moment of the rotor could be overcome by employing two unhinged rotors rotating in opposite directions.

As an extension of this method, US Pat. No. 3,409
No. 248 selectively changes the differential lateral periodic pitch as a function of the helicopter's forward speed to offset or reduce roll moments and optimally direct the lift vector of each rotor. In addition, a method for obtaining the optimum lift-drag ratio characteristics has been suggested.

In order to realize this function, the above-mentioned U.S. Patent No. 3,409,248
The mechanism disclosed in that issue was a simple linkage that provided direct input to the control rods of each rotor, either manually or by means of an air velocity detector using a computer to obtain the correct gain.

U.S. Pat. No. 3,570,786 discloses that the differential lateral periodic inputs as a function of human power from the simultaneous control levers are controlled by coupling the simultaneous control levers to the differential lateral periodic inputs of the control system. The Buchi method of obtaining pitch has been suggested.

According to US Pat. No. 3,570,786, a constant differential lateral periodic input provides adequate efficiency in high speed flight where the simultaneous control lever position is constant.

U.S. Pat. No. 3,521,971 describes a gyroscopic system applied to coaxial twin-rotating hingeless rotors during helicopter motion.
It has been recognized that moments cause bending stresses and deflections of the vane in opposite directions.

By increasing the fixed phase angle and introducing a differential periodic pitch for each rotor, aerodynamic moments are created that offset the gyroscopic moments due to precession during these flight motions.

Therefore, the above-mentioned U.S. Pat. No. 3,409,249 teaches that if a differential lateral periodic pitch is introduced as a function of the forward speed of the helicopter in a coaxial twin-rotating hingeless rotor, the lift of the rotor can be adjusted to a constant 1H. It has been suggested that the optimal lift-drag ratio characteristics of the rotor blade in steady state can be obtained by orienting the rotor directly.

The above-mentioned US Pat. No. 3,409,249 itself does not mention the cancellation of gyroscopic moments by precession.

Gyro moment I due to precession that occurs during movement
In order to compensate for
Known by No. 71.

However, the present inventors have come to the realization that a large fixed phase angle can only compensate for precession moments at one flight speed.

The present invention automatically orients the lift vectors to the rotor in a twin-rotating hingeless rotor blade in an optimal manner to obtain the optimal lift including canceling the roll moment by the lift vectors in opposite directions. - Obtain the drag ratio and
The objective is to automatically cancel or minimize precessional gyroscopic moments at all flight speeds, thereby achieving optimal performance and maneuverability of the helicopter in both steady state and motion conditions.

In accordance with the present invention, as a control method for a twin-rotating hingeless rotor, the phase angle of each rotor is changed as a function of the forward speed of the helicopter, so that a selected differential periodic pitch can be applied to the rotor. periodic pitch generated in response to human forces to selectively direct the rotor's lift vector to obtain an optimal lift-drag ratio, and also due to the precession of the rotor at all flight speeds. Offset or minimize gyroscopic moments.

In accordance with a preferred embodiment of the present invention, the phase angle of each rotor is controlled during flight through a phase angle range between about 20 degrees and 70 degrees for flight speeds between hover and 160 knots. -Variable as a function of speed.

Furthermore, according to the invention, the phase angle of the rotor is automatically changed during flight as a function of forward speed, and also serves to link the longitudinal periodic input to the transverse periodic control, so that The desired pitch control input produces the desired differential periodic pitch between the two rotors.

A preferred means for changing the phase angle of each rotor blade is an analog mixer based on a rotating swash plate mechanism, the rotating swash plate being able to tilt around an arbitrarily selected tilt axis; It is also connected to the swash plate of the helicopter rotor blade so as to tilt the swash plate of the rotor blade around a periodic pitch control tilt axis in synchronization with the rotor blade.

Here, periodic pitch control is provided to the main rotary swashplate through an analog mixer, most importantly because the tilt axis of the analog mixer can be selectively oriented as a function of helicopter forward speed. helicopter
Modification of the rotor phase angle as a function of forward speed is accomplished without providing periodic control inputs from the analog mixer to the rotor swashplate.

In accordance with the present disclosure, an analog mixer is used in a helicopter rotor control system to operate and tilt about a selected tilt axis in response to periodic pitch inputs from a control lever. providing a periodic control input to the rotor by creating a similar tilt in the rotor's swashplate, which also serves to vary the phase angle of the rotor during flight as a function of helicopter speed; It has the ability to perform this phase angle changing function without applying periodic control inputs to the rotor.

Referring to FIG. 1, a coaxial twin-rotating unhinged rotor and its control system is further indicated by the reference numeral 10, the rigid rotors 12 and 14 being supported by a helicopter fuselage in a conventional manner. , generate lift by rotating about a common axis 18.

Note that the rotary blades 12 and 14 are shown side by side for convenience of illustration.

Each rotor has its own control system, preferably the same surface control system, and includes a conventional swashplate assembly 20 consisting of a fixed swashplate section 20a and a rotating swashplate section 20b. There is.

The swash plate portion 20b is connected to the pitch blades of the rotor blades 12 and 14 by a conventional linkage;
Movement of swashplate 20 along rotor axis 18 causes simultaneous pitch changes in the rotor blades, and inclination of swashplate 20 with respect to axis 18 causes periodic pitch changes in the blades.

The rotating swash plate assembly 20 is translated and/or tilted by a main servo mechanism 22.

The servo mechanism 22 is controlled by inputs from the simultaneous control lever 24, the periodic control lever 26, and the foot pedal 30, and these input forces are applied to the mixer mechanism 2.
8 to the servo mechanism 22.

Note that the input from the lever 26 passes through an analog mixer 40 before entering the mixer mechanism 28.

For purposes of describing the invention, it will be sufficient to describe the rotor control system in the schematic diagram shown in FIG.

The rotor blade and its control system 10 are disclosed in U.S. Pat. No. 340
No. 9249, No. 3521971 and No. 3570
In operation, a periodic control lever 26 is used to control the pitch and roll of the helicopter, as described in more detail in that patent.

Simultaneous control lever 24 is used for vertical control, and pedal 30 is used for directional or yaw control.

The simultaneous control lever 24 uniformly changes the blade pitch angle of each rotor blade for rotor blade thrust control.

The periodic control lever 26 periodically and uniformly changes the blade angle of each rotor to control the pitch and roll moment of the rotor.

A periodic input in the longitudinal direction produces a pitch motion, and a periodic input in the lateral direction produces a roll motion.

Foot pedals 30 vary the blade angle of each rotor simultaneously, but equally in opposite directions, for directional or yaw control.

That is, when the simultaneous pitch of the upper rotor 14 increases and that of the lower rotor 12 decreases due to the action of the pedal 30, the torque generated by the upper rotor 14 in the counterclockwise direction increases, causing the nose to turn to the left. A yawing moment 1~ is added to the aircraft.

The effect of pedal 30 increases the simultaneous pitch of lower rotor 12 and decreases the simultaneous pitch of upper rotor 14, creating a moment that yaws the nose to the right.

For purposes of the following discussion, rotor 14 is the upper rotor and rotates counterclockwise when viewed from above, and rotor 12
is the lower rotor blade, and it is assumed that it rotates clockwise when viewed from above.

Helicopters do not respond immediately to feathering changes by the pilot.

The development of maximum blade bending moments and helicopter control moments is delayed relative to changes in blade pitch angle due to the flexibility of the rotor blades 12 and 14.

The meaning of this delay can best be understood from FIG.

Let's now assume that we want to apply a pitching moment to the nose of the helicopter at an azimuth angle of 180 degrees, and change the pitch of the blades when the blades pass a position at an azimuth angle of 180 degrees.

Due to the delay in the generation of the control moment described above, the helicopter does not experience the maximum control moment due to changes in its blade pitch at the nose or 180° azimuth position, but rather at an angle from there in the direction of blade rotation. For example, at a position of 180°+θ, that is, at an azimuthal position M in FIG.

As can be seen from FIG. 2, as a result of the delay of the angle θ, the forward (longitudinal) displacement of the periodic control lever 26 causes pitching moment (MP) and roll moment) (MR) of each rotor blade. Both components will be produced.

The pitching moment Mp is simultaneous and also produces the desired pitching acceleration in the helicopter.

Since the rotors 12 and 14 rotate in opposite directions, the roll moment components MR caused by the longitudinal movement of the periodic control lever cancel out, but these components create undesirable bending stresses in the axis of each rotor. .

The retardation angle θ is approximately 30° for this type of rotor blade.

Bending stresses between these axes are minimized by adjusting the swash plate 20 of each rotor to compensate for the control moment delay.

If the swash plate 20 of each rotor blade is deflected in the direction opposite to the rotation of the rotor blade by an amount equal to the delay angle θ, as shown in FIG. 180
The bending stress between the axes will be reduced, occurring only at an azimuth angle of .

As is clear from Figure 3, pitching moment) M
Only p is the nose of the helicopter (180° azimuth position)
are generated on the rotor blades 12 and 14 along a longitudinal or longitudinal line passing through the rotor blades 12 and 14.

As illustrated in FIG. 3, changing the pitch does not create a roll moment component MR.

However, as a result of the pitch acceleration of the helicopter, the gyroscopic precession characteristics of the rotor create an additional moment designated QMG.

The tendency of the moment of the upper rotor to cause roll is counteracted by that of the lower rotor, but on the other hand creates undesirable blade bending stresses and deflections.

As shown in U.S. Pat. No. 3,521,971, if the rotating swashplate 20 is exposed to a significantly large angle, e.g., approximately twice the phase angle gamma (F) as shown in FIGS. If adjusted, Figures 4 and 5
The results shown in the figure are obtained.

That is, an aerodynamic moment is automatically generated that balances the unfavorable blade bending moment due to gyroscopic moments caused by changes in helicopter attitude during motion.

In this way, overstressing of the rotor blades is avoided and
Additionally, the necessary blade tip clearance is maintained during movement.

FIG. 4 illustrates the effect of applying a controlled pitch change input at azimuth position ΔP to produce an aerodynamic blade bending moment 1~ at azimuth position M of rotor blades 12 and 14. There is.

The resulting moment)-M in FIG. 4 is the pitch change moment) Mp which causes the blade to pitch, and the transverse aerodynamic moment MA, which corresponds to the roll moment l''MR shown in FIG. It can be decomposed into vectors.

What should be noted in FIG. 4 is that the helicopter is in motion, ie, changing attitude in response to a moment applied to the helicopter.

As fully explained in U.S. Pat. No. 3,521,971, a characteristic of a helicopter rotor is that a gyroscopic moment is applied to the helicopter at an azimuth position G 90 degrees forward of the attitude change moment M in response to an attitude change. act.

The gyroscopic moment shown at location G is combined with the aerodynamically applied pitching moment l-Mp to produce an additive gyroscopic moment 1-GP that causes the helicopter to pitch. is Peltle decomposed into a lateral gyroscopic moment MG corresponding to a roll moment MR.

As can be seen from FIG. 4, the roll moment (MA) aerodynamically generated in each of the rotor blades 12 and 14 cancels the gyroscopic moment M6, so that the blades due to the gyroscopic force generated during the motion of the helicopter The bending moments cancel out.

That is, FIG. 4 shows that by proper selection of the phase angle gamma, a differential pitch coupling resulting in an aerodynamic moment MA is automatically generated on each rotor blade, which in turn generates a gyroscopic coupling occurring during the motion of the helicopter.・Momen) This shows that it cancels out the reverse effect caused by Mc.

It will be clear to those skilled in the art that the aerodynamic moment MA of FIG. 4 always subtracts the effect of the gyroscopic moment M6 and also cancels M6 when MA''M6.

FIG. 5 illustrates that advantages are automatically obtained by selecting the phase angle gamma equal to the phase angle of FIG. 4 even during steady state flight rather than during helicopter motion.

Looking at FIG. 5, where in steady state flight the phase angle gamma is again chosen to be substantially larger than the delay angle θ, when the control input causes a blade pitch change at position ΔP, the differential pitch coupling is An aerodynamic moment l''MA is generated on each rotor blade at .

This aerodynamic moment is equal to the additive pitch moment (MP) and the countervailing roll moment (MR).
can be decomposed.

The generated roll moment) MR is the rotor blade 12 and]
4 to selectively direct the lift vector to each rotor blade 12 and 14, thus achieving the optimal
Creates a lift-drag ratio to optimize rotor efficiency.

Therefore, as can be seen from Figures 4 and 5,
By utilizing the selected phase angle gamma, the helicopter
During the motion of the aircraft, an aerodynamic moment MA is generated which cancels out or reduces the gyroscopic moments 1 to M6 caused by attitude changes of the helicopter or rotor during the motion, and during steady state flight, the aerodynamic moment MA is lateral periodic pitch to optimally orient the lift vector of each rotor, resulting in an optimal lift-to-drag ratio for the rotor.

The significance of the cancellation of the gyro moment M6 explained with reference to FIG. 4 can be best understood from FIG.

FIG. 6 is a front view of a helicopter using coaxial twin-rotating type unhinged rotor blades, in which the gyro moment (MG) moves the rotor blades 12 and 14 from the solid line to the position where the gyro force is applied as shown by the dotted line. , and therefore,
The tip of the blade A of the rotor blade 120 and the tip of the blade R of the rotor blade 14 are close to each other.

By automatically introducing differential lateral pitch coupling as a function of longitudinal pitch input, an aerodynamic moment is created that is equal and opposite to the gyroscopic moments 1 to M6 caused by precession. be done.

This action keeps the two rotor blades in the position shown by the solid line in FIG. 6, eliminating the problem of tip clearance and minimizing the stress applied between the blades and the shaft.

Figure 6 illustrates that the offset lift vector that produces the optimal lift-drag ratio characteristics produces a tilt of the rotor blade as shown by the solid line □, and this is particularly important for maintaining appropriate tip clearance at high speeds. This is why gyro moment cancellation is so important.

Although U.S. Pat. No. 3,521,971 discloses the use of a fixed phase angle to offset the gyroscopic moments created during helicopter motion, the inventor of the present application has determined that the phase angle can be adjusted during flight to compensate for the helicopter's forward speed. V as a function of V to obtain optimal lift-drag ratio and stall margin at all speeds, and to offset gyroscopic moments especially at high speeds, as well as differential lateral periodic control. We have found that this is preferable because it eliminates the need for separate input.

This need is best understood from FIGS. 7, 8, and 9.

Figures 7 to 9 show that for a given motion of the helicopter, e.g. a motion of l rad/s2, as the forward speed V of the helicopter changes, the gyroscopic moments occurring during this motion are offset. Alternatively, the need to use various rotor phase angles to generate the aerodynamic moment to be minimized is illustrated.

First, let us consider the motion of a helicopter when it is flying at 150 knots 1-(k).

As shown in FIG. 7, vertical sequential control input A
Pitching moment or pitching acceleration Mp (150k) for 1 (150k) to start motion
is necessary for the occurrence of

Looking at Figure 8, the pitching moment l=Mp (150k) obtained in Figure 7 is the gyro moment M.

(150k).

Now looking at Figure 9, we see that when a large phase angle F1 is used, the same longitudinal periodic input A as in Figure 7, (15
0k) yields an aerodynamic moment)MA(150k) which is equal to the gyroscopic moment Mc(150k).

Now assume that the helicopter is flying at a forward speed of 50 knots instead of 150 knots, and that the pilot decides to make the same 1 rad/S2 motion as before.

At a forward speed of 50 knots, a larger longitudinal periodic pitch A, (5Qk) must be applied to produce this motion.

Figure 7 shows this A, (the same pitching moment MP (50k) as at 150 knots due to 5010)
It has been shown that this occurs.

It will be appreciated that MP(150k) is equal to MP(50k) since the same motion is being performed.

Now, proceeding to Figure 8, pitching moment) Mp (
50k) yields the gyro moment I.MG (50k), which is equal to Ma (150k).

Looking at Figure 9, when applying the same longitudinal periodic pitch input A, (50k) as in Figure 7, the aerodynamic moment )-MA ( 50k), it is found that it is necessary to use a smaller phase angle P3.

If the phase angle were kept at Pl during both the 150-knot motion and the 50-knot motion, an aerodynamic moment significantly larger than the gyroscopic moment would occur during the 50-knot motion, and both moments would be It is clear from FIG. 9 that no cancellation occurs.

The inventor of the present application has found that, as a characteristic of a helicopter using this type of rotary blade system, the degree of change in longitudinal periodic input is
It was discovered that it was necessary to overcome the drag forces and keep the helicopter in equilibrium at all forward speeds.

It is known from U.S. Pat. No. 3,409,248 that the magnitude of the differential lateral periodic pitch input required to produce an optimal lift vector offset is substantially as represented in FIG.

Through analysis and development of helicopters propelled by coaxial twin-rotating unhinged rotors, the inventors have determined that the phase angle P of rotors 12 and 14 is adjusted to It was determined that the speed range needed to be varied to obtain optimal performance and maneuverability throughout the speed range.

The curves in FIG. 11 indicate desirable phase angles in three speed ranges.

That is, for flight speeds from hovering to about 80 knots, a phase angle of about 20° is used;
For flight speeds from 0 knots I to about 140 knots, the phase angle is gradually increased from about 20° to about 70°, and for flight speeds from about 140 knots to about 160 knots I. A phase angle of approximately 70° is used.

Although FIG. 11 shows preferred phase angle changes in this case, it is clear that the phase angles may also be changed on both sides.

The curve in Figure 11 is designed to obtain the above results and characteristics.
This is the preferred curve in this case, taking vibration and balance into account.

This change in phase angle as a function of helicopter forward speed acts to combine the longitudinal periodic pitch and the lateral periodic pitch, as explained in FIG. Applying the longitudinal periodic pitch input (shown in FIG. 10) necessary to keep the winning helicopter in equilibrium at all forward flight speeds causes the rotors 12 and 14 to have a differential lateral periodic pitch. 1. Occurs dynamically.

This differential lateral periodic pitch is determined according to the desired curve of FIG.
By creating a lift vector I, an optimum lift-to-drag ratio characteristic can be obtained, thereby optimizing the efficiency and maneuverability of the rotors 12 and 14.

In addition, as shown in FIG. 4, gyroscopic moment cancellation or reduction occurs as described above.

Preferably, the rotor phase angle is varied during flight as a function of helicopter speed by an analog mixer 40, shown schematically in FIG. 1 associated with each rotor's swashplate 20.

Since the same analog mixer is used for the rotors 12 and 14, only the analog mixer for the rotor 14 will be shown in detail in FIGS. 13 and 14.

Analog mixer 40 is coupled to fixed swashplate portion 20a of rotary swashplate assembly 20, and rotary swashplate portion 20b is supported by the fixed swashplate portion for rotation about axis 41.

This rotary swash plate portion 20b is connected to the push rod member 44 in a normal manner.
is connected to the pitch 46 of each blade 50 of the rotor blade 14.

Translation of the rotating swash plate 20 with respect to the axis 41 causes a simultaneous pitch change in all vanes 50.

On the other hand, pivoting the swashplate 20 about any horizontal axis that intersects the axis 41 causes periodic pitch changes in the blades 50.

Analog mixer 40 is coupled to rotary swashplate 20 by establishing a horizontal axis as rotary swashplate 20 tilts with longitudinal and lateral cyclic control inputs from cyclic control lever 26. , rotor phase angle gamma (
I').

A mixing unit 28 of conventional design for combining periodic and simultaneous inputs is preferably arranged between the analog mixer 40 and the rotary swashplate 20, so that the periodic control output of the mixer 40 is arbitrary. is applied to the rotating swashplate 20 with the necessary coupling, gain, etc., to produce the required change in phase angle as a function of said speed.

The upper non-rotating member 56 of the analog mixer 40 is engaged by an engagement member 58 to the fuselage or other parts so as not to rotate about the mixer axis 42.

As best seen in FIG. 14, the lower rotatable plate member 60 is connected to the non-rotating member 56 by a low-friction internal bearing 62, so that the lower rotatable plate member 60 rotates about the axis 42 relative to the non-rotating member 56. Can be rotated.

This lower plate member 60 is connected to runners 1 to 64 of a universal joint member 66, and this universal joint member 66 has a low friction so as to be able to freely rotate around the (preferably vertical) mixer axis 42. It is supported by the fuselage body and other fixed members via bearings 68.

Universal joint 66 allows members 56 and 60 to pivot in unison about any horizontal axis intersecting axis 42 as commanded by control inputs.

Furthermore, when the universal joint 66 rotates around the axis 42,
Plate 60 undergoes rotation about axis 42 relative to plate 56.

These plate members 56 and 60 are tilted or pivoted about a selected horizontal axis by control input movements applied by a lateral periodic pitch input member 70 or a longitudinal periodic input member 72 from a steering lever. arise.

Members 70 and 72 pivot bellcranks 74 and 76 about axes 78 and 80, thereby providing input rods 82 or 84 with lateral or longitudinal cyclic control input movement to rotating swashplate members 56 and 60. , so that these swashplate members pivot or tilt about a selected horizontal axis.

The input rods 82 and 84 are connected to the bell crank 7 via ball bearings.
4 and 76 and the rotating swash plate member 60.

Rods 82 and 84 are coupled to swashplate member 60 at azimuthal positions spaced 90 degrees apart.

. The non-rotating member 56 is connected to the mixer unit 28 in a conventional manner.
is coupled to the main swashplate non-rotating member 20a through the analog mixer swashplate members 56 and 6.
Performing a tilting motion around the tilting axis where 0 is selected,
It is accurately transmitted to the main rotary swashplate 20 to establish the selected periodic pitch change of the vanes 50.

Analog mixer member 40 is coupled to swashplate 20 via mixer unit 28 by rod members 8'6, 88 and 90, as schematically shown in FIG.

In operation, the analog mixer 40 is connected to the drive mechanism 9.
2 causes the phase angle of the rotor 14 to vary as a function of the flight speed V of the helicopter.

That is, in response to a command from the flight speed detection mechanism 94, the drive mechanism 92 operates the universal joint 66 and therefore also the rotating swash plate 60.
, relative to the swashplate member 56 about the axis 42, thereby causing the swashplates 56-60, and thus also the main swashplate 20, to rotate through the control rods 70 and 72 from the steering levers. Establishing a horizontal axis for tilting in response to periodic inputs.

This selection of the horizontal axis during tilting of the main swashplate determines the azimuthal position ΔP of the rotor at which the blade pitch change occurs as a result of the input from the periodic control lever, and therefore also determines the phase of the rotor. The angle gamma (I') is determined.

By driving the drive mechanism 92 with a predetermined functional relationship in response to commands from the flight speed detection mechanism 94, the rotor phase angle will change as a function of the helicopter's forward speed V during flight.

As an important disclosure of the present invention, bell cranks 74 and 7
Since the working radius of 6 coincides with the axis 42 of the analog mixer, the member 60 by the flight speed detection mechanisms 92-94
The phase change rotation causes little displacement of the control rods 82 and 84 along the surface of the cone, thereby avoiding the application of undesirable control inputs to the rotor during phase change rotation of the analog mixer 40. can.

Although analog mixer 40 is shown in this application as being used with an unhinged rotor, it will be apparent to those skilled in the art that it may be similarly applied with hinged or other types of rotors capable of varying phase angles. Will.

Although the present invention has been described in connection with a coaxial twin-rotating hingeless rotor for a helicopter, it may also be applied to any twin-rotating hingeless rotor system and rotor used outside the helicopter field. It will be clear to those skilled in the art that similar applications can be made.

Throughout the above discussion, we have talked about moment cancellation; however, complete moment cancellation can only be achieved under special conditions; under other conditions, partial moment cancellation is required to minimize the effects of unfavorable moments. It will be clear to those skilled in the art that this can be achieved.

The scope of the invention is not limited to the details of construction shown and described above, and various modifications may be made by those skilled in the art.

[Brief explanation of drawings]

FIG. 1 is a schematic diagram of a control system according to the invention, with two rotors shown side by side for convenience of illustration and explanation. FIG. 2 shows the effect of delay angle on the response to longitudinal inputs at the upper and lower rotors in an uncompensated condition. FIG. 3 shows a state in which the influence of the delay angle has been compensated for. This figure also shows the vector of the gyroscopic moment caused by the pitch acceleration. FIG. 4 illustrates a vector relationship that uses phase angle to generate aerodynamic moments in the upper and lower rotors, thereby canceling out or reducing the gyroscopic moments that occur during helicopter motion. Figure 5 shows that during steady-state flight, the phase angle is used to generate a differential lateral periodic pitch between the two expansions to offset roll moments and optimize lift. The vector/le relationship that directs the lift vector of each rotor to obtain the drag ratio characteristics is shown. Figure 6 is a front view of a helicopter using coaxial twin-rotating hingeless rotors, showing the effect of the gyroscopic moment on the rotor during motion, and the gyroscopic moment due to the generation of aerodynamic moments. The results are shown after stabilizing by compensating for I. FIG. 7 is a graph showing the relationship between longitudinal periodic pitch A1 and pitching moment Mp for various helicopter forward speeds. FIG. 8 is a graph showing the relationship between pitching moment Mp and gyro moment M. FIG. 9 is a graph showing the relationship between longitudinal periodic pitch A1 and aerodynamic moment 1"MA for various rotor phase angles. FIG. 10 shows a coaxial twin-rotating hingeless rotor. At forward flight speeds between hover and 150 knots, the helicopter can overcome drag and
3 is a graph showing the longitudinal periodic pitch A1 required to provide a stable flight attitude to the aircraft. FIG. 11 is a graph showing the rotor phase angle required to obtain optimal rotor characteristics and maneuverability between hover and 160 knots without adding separate differential lateral periodic pitches. be. Figure 12 shows how to hover at 150 knots, including canceling roll moments and selectively directing the lift vector I to obtain an optimal lift-drag ratio during steady-state flight. 2 is a graph showing the differential lateral periodic pitch input required to obtain optimal characteristics and maneuverability between FIG. 13 is a schematic perspective view of an analog mixer used to vary rotor phase angle as a function of helicopter forward speed. FIG. 14 is a sectional view of the universal joint portion of the analog mixer of FIG. 13. 10~Rotor blade and control system (collective), 12.14
~Rotor blade, 18~Rotor blade axis, 20~Rotary swash plate, 203
~Same fixed part, 20b~Same rotating part, 22~servo mechanism, 24~simultaneous control lever, 26~periodic control lever, 28~mixer mechanism, 40~analog mixer, 41~
Axis of main rotating swash plate, 42 - Axis of analog mixer rotating swash plate, 44 - Push rod, 46 - Pitch, 50 - Feather blade 6 - Non-rotating member of rotating swash plate, 60 - Rotating plate member ,
58~Engagement member, 62~Bearing, 64~I/Runion, 6
6 ~ Universal joint, 68 ~ Bearing, 70 ~ Horizontal periodic pitch input member, 72 ~ Vertical direction periodic input member, 74.76 ~
Bell crank, 78° 80 ~ coaxial line, 82.84 ~ input rod, 86, 88.90 ~ rod member, 92 ~ drive mechanism, 94
~Flight speed detection mechanism.

Claims (1)

[Claims] 1. In a method for controlling an aircraft having twin-rotating hingeless rotor blades, the phase angle gamma of each rotor blade is increased by an equal angle in opposite directions as the forward speed of the aircraft increases. as a function of the forward speed of the aircraft. 2. In the method of claim 1, the phase angle gamma is varied between about 20° in hovering and about 70° at a flight speed of about 150 knots. How to characterize it. 3. In the method of claim 1, the phase angle gamma is approximately 20 degrees at flight speeds from hovering to approximately 80 knots, and approximately 20 degrees at flight speeds from approximately 80 knots to approximately 140 knots. between about 20° and about 70°, and to about 70° at flight speeds from about 140 knots to about 160 knots. 4. A rotary swash plate assembly 20 having bi-rotating hingeless rotor blades and tiltable about an axis to periodically control blade pitch for each rotor blade; and the rotary swash plate assembly. an analog mixer 40 connected to the aircraft, and further comprising means for varying the angular position of the rotary swashplate about said axis. 5. In the aircraft of claim 4, the analog mixer 40 is arranged concentrically around a first axis 42 and is held against rotation, but intersecting the first axis 42. and a first member 5 which performs a tilting movement about all tilting axes lying in one plane perpendicular to the second axis.
6, a second member 60 that tilts together with the first member and rotates around the second axis; and an action that connects the first member to the swash plate and tilts the swash plate. means 86, 88, 90 for applying a tilting force to said second member 60 in a first azimuthal position, thereby causing a tilting force about a first of said tilting axes. first means 70 for producing a tilting movement of said first and second members; 74. 82 and 9 from the first azimuth position.
said second member 6 at a second azimuthal position spaced apart by 0°;
second means 72.7 for applying a tilting force to 0, thereby tilting said first and second members about a second of said tilting axes perpendicular to said first tilting axis;
6.84 and the second and second azimuth angles connected to the second member 60 and preventing the second member from rotating about the first axis 42 in response to flight speed; means for selecting a position 66, 92. An aircraft characterized by being provided with 94. 6. In the aircraft of claim 5, said first and second means each have a first end connected to said second member 60 at one of said azimuthal positions. Bar member 82
．． 84 and a bell crank 74, 76 connected to the other end of the rod member, the operating radius of the control force input side of the bell crank being aligned with the first axis 42. An aircraft featuring