Rotary wing aircrafts

Rotary wing aircrafts comprise a fuselage having a rotor driving mast extending substantially vertically. A rotor comprising a plurality of rotors is mounted on the mast for at least feathering and flapping movement. The flapping hinge is offset radially outwardly from the axis of rotation. Further, the rotor blades have weights mounted on the outboard tip ends thereof. The arrangement provides a significantly improved controllability of the aircraft.

The present invention relates to rotary wing aircrafts and more 
particularly to rotary wing aircrafts having rotors tiltable with respect 
to rotor driving masts. 
Conventional rotary wing aircrafts include rotors which provide lifts for 
supporting the aircrafts in air. In most of such rotary wing aircrafts, 
the rotors are also relied on in producing horizontal thrusts for 
initiating and sustaining horizontal movement of the aircrafts. 
In conventional rotary wing aircrafts, the rotors are mounted, with a very 
small number of exceptions, in such a manner that it can be tilted with 
respect to rotor driving masts. For example, in a so-called fully 
articulated rotor, rotor blades are mounted on a rotor hub for movements 
about a flapping and a dragging axes as well as about a feathering axis. 
In operation, each blade assumes a position wherein weight, centrifugal 
force, aeronautical force and inertia force acting thereon balance each 
other. When it is desired to produce a horizontal movement, the pitch or 
feathering angle of the rotor blade is cyclically changed so as to cause a 
cyclic change in the lift produced by the blade. This causes a cyclic 
change in flapping angle of the rotor so that the plane of rotor rotation 
is in effect changed. In a so-called "see-saw" type two blade rotor 
system, the rotor plane can similarly be tilted with respect to the rotor 
mast through a cyclic pitch control. 
In these known types of rotary wing aircrafts, problems have been 
experienced in that the aircraft shows only a slow response to an 
actuation of a control device. For example, when the control device is so 
actuated that a forward thrust is produced, the cyclic pitch control 
mechanism is immediately actuated in a desired direction by a desired 
amount. Then, a cyclic change is produced in the lift on each rotor blade 
so as to produce a moment for causing tilting of the rotor plane with 
respect to the rotor mast. Thus, a forward thrust component is produced in 
the rotor and the aircraft starts to advance. Thereafter, the aircraft 
fuselage conducts a nose-down movement to some extent to reduce the tilt 
angle between the rotor plane and the mast. In this manner, it takes a 
substantial time before the aircraft follows an actuation of the control 
device. 
Further, when the aircraft is in a steady state and the rotor plane is 
tilted under a turbulence with respect to the rotor mast, such tilt cannot 
be sensed until the aircraft has started to move from the steady state. 
Any corrective action for such turbulence can be initiated only after such 
movement has been sensed. 
Another problem in these types of rotors is that although the control 
device, such as a control stick positively determines the position of the 
cyclic pitch control mechanism, the tilt angle of the rotor plane with 
respect to the rotor mast cannot be determined only by the control device 
but changes in accordance with a change in attitude of the aircraft. 
From the above reasons, control of rotary wing aircrafts has been 
considered very difficult. It has been believed that a time-consuming 
training is required for training pilots for helicopters. 
The present invention has therefore an object to provide a rotary wing 
aircraft which has one or more tiltable rotors but is easy to control. 
Another object of the present invention is to provide a rotary wing 
aircraft which has an ability of rapid response to a control operation. 
A further object of the present invention is to provide a rotary wing 
aircraft which has tiltable rotor means but in which the tilt angle 
between the center axis of the rotor plane and the rotor mast can be 
maintained within a very small range. 
Still further object of the present invention is to provide a rotary wing 
aircraft which has a good controllability as compared with conventional 
rotary wing aircrafts. 
Theoretically, there are various factors which affects the controllability 
of the rotary wing aircraft. Those are the weight (W) of the aircraft 
under a standard operating condition, the height (h) or vertical distance 
between the center of gravity and the center of the rotor plane, the 
moment of inertia (Io) of the aircraft about the pitching axis passing 
through the center of gravity, the number of rotor blades (b), the rated 
operating angular velocity (.OMEGA.o) of the rotor, the centrifugal force 
(Zo) acting on the rotor blade at the rated operating angular velocity 
(.OMEGA.o), the offset distance (e) of the flapping hinge from the 
rotating axis of the rotor, the mean chord length (C) of the rotor blade, 
the moment of inertia (I.sub.R) of the rotor blade about the flapping 
hinge, the radius (R) of the rotor, and the density of the air (.rho.o) 
under the standard atmospheric condition. 
The inventor has found that the values 
##EQU1## 
have very important effects on the controllability. In defining the value 
T, it has been assumed that the ratio of the lift coefficient of the rotor 
blade section to the attack angle is 5.73. 
More specifically, the ratio 2.sqroot.K/T is of particular importance. With 
the ratio greater than one, the tilt angle between the rotor plane and the 
rotor mast can be maintained very small. This means that the aircraft 
fuselage moves very quickly in response to the tilting movement of the 
rotor plane so that the controllability of the aircraft is remarkably 
improved. Further, when a helicopter is subjected to a sudden gust in 
hovering, the rotor plane may be tilted with respect to the rotor mast 
under the influence of such gust, however, such tilting movement can be 
immediately sensed by the pilot and corrective action can be quickly 
taken. According to a further feature of the present invention, the 
ultimate angular velocity of the fuselage which is attained by an 
actuation of the control member such as a swash plate for a unit angle is 
much smaller than that in conventional helicopters, so that it is possible 
to actuate the control member by a greater angle than in the conventional 
helicopters without encountering any danger. This will provide an 
increased safety to the aircraft. Practically, with the ratio greater than 
0.8, a significant improvement on the controllability can be accomplished. 
In attaining a larger value of the ratio, the design factors may be 
determined in such a manner that the item K is increased and/or the item T 
is decreased. Among the factors which affect the value K, the weight W and 
the height h are not versatile factors and they must be determined from 
other design requirements. Similarly, the blade number 6 may be, in large 
part, determined from the aeronautical or other design requirements. Thus, 
the centrifugal force Zo and the offset value e must be properly increased 
in order to increase the value K. For example, the rotor blades may be 
provided with weights at or in the vicinity of tip ends thereof to obtain 
increased value of the centrifugal force. Further, the flapping hinge of 
each rotor blade may be located as far as possible from the rotating axis 
of the rotor unless there is any objectionable reason. 
Among the factors affecting the value T, the density of the air .rho.o 
should be constant. The chord length C of the rotor blade, the rated 
angular velocity .OMEGA.o and the radius R of the rotor may be, in large 
part, determined from other design requirements. Therefore, the most 
effective way for decreasing the value T is to increase the moment of 
inertia I.sub.R of the rotor about the flapping hinge. A weight provided 
at or in the vicinity of the rotor blade will be effective to increase the 
moment of inertia I.sub.R and therefore to decrease the value T. 
According to a preferable mode of the present invention, each rotor blade 
is made of a one-piece construction of an aluminum alloy extrusion. This 
structure provides an increased centrifugal force and an increased moment 
of inertia of the rotor blade about the flapping hinge so that the 
aforementioned ratio 2.sqroot.K/T is increased to a satisfactory level. 
Further, the one-piece construction of the aluminum alloy extrusion 
provides an adequate strength to withstand the aeronautical and mechanical 
forces to which the blade is subjected in use. 
In installing a weight on the rotor blade, the weight may not be secured to 
the rotor blade but simply fitted thereto and maintained in position by a 
cable or other tension member extending lengthwise in the blade. This 
arrangement will be advantageous in that the blade itself is not subjected 
to the centrifugal force acting on the weight and consequently the blade 
can be of thinner wall construction.

Referring now to the drawings, particularly to FIG. 1, there is 
schematically shown a helicopter embodying the feature of the present 
invention. The helicopter comprises a fuselage 1 having an engine 2 
mounted thereon. The helicopter 1 further includes a vertically extending 
mast 3 which is drivingly connected with the engine 2 in a manner well 
known in the art. The fuselage 1 has a tail boom 4 extending rearwardly 
therefrom and being equipped with a tail rotor 5. As well known in the 
art, the tail rotor 5 is also driven by the engine 2. 
At the top end of the mast 3, there is secured a rotor hub 6 to which a 
plurality of rotor blades 7 are installed. In the particular embodiment 
shown in FIG. 1, three rotor blades 7 are installed on the rotor hub 6 at 
angularly equi-distant positions. Each of the rotors 7 is mounted on the 
rotor hub 6 for flapping movement about a flapping axis 8 which is 
provided by a flapping hinge pin and which is, according to one of the 
features of the present invention, offset radially outwardly by an offset 
distance e. The blade 7 is of course mounted on the rotor hub 6 for 
feathering movement by means of a known mechanism about a feathering axis 
which extends longitudinally through the blade 7. 
The rotor mast 3 and the hub 6 is rotated by means of the engine 2 about a 
longitudinal axis 3a of the mast 3 and consequently the rotor rotates with 
a radius R at the angular velocity .OMEGA.o. In the illustrated example of 
the rotor blade 7, a weight 9 is installed at the tip end thereof. The 
weight 9 is not secured to the rotor blade 7 but simply fitted thereto and 
connected to the flapping hinge pin by means of a wire 10 extending 
through the blade 7. 
The helicoper has a center of gravity G located substantially on an 
extension of the longitudinal axis 3a of the mast 3, and the rotor hub 6 
which has a center 6a located with a vertical distance h from the center 
of gravity G. 
As previously described, the controllability of a helicopter is affected by 
the physical values 
##EQU2## 
wherein Io is the moment of inertia of the helicopter about the pitching 
axis passing through the center of gravity, b the number of rotor blades, 
.OMEGA.o the rated operating angular velocity of the rotor, Zo the 
centrifugal force acting on the rotor blade at the rated operating angular 
velocity, C the mean chord length of the rotor blade, I.sub.R the moment 
of inertia of the rotor blade about the flapping hinge, and .rho.o the 
density of air under the standard atmospheric condition. 
In the embodiment shown in FIG. 1, the value 2.sqroot.K/T is determined as 
being greater than 0.8, more preferably greater than 1 by locating the 
flapping hinge 8 at a position offset from the axis of rotation 3a and 
mounting the weight 9 at the tip of each rotor blade 7. Further, as shown 
in FIG. 2, the rotor blade 7 is made of an aluminum alloy extrusion. This 
is effective to increase the moment of inertia I.sub.R of the rotor blade 
7 about the flapping hinge 8 and consequently decrease the value T. 
Practically, it is preferred that the value K is not smaller than 25 
1/sec.sup.2 and the value T is not larger than 12 1/sec. 
In order to show the advantageous features of the present invention, 
theoretical descriptions will now be made taking reference to FIG. 3. In 
FIG. 3, the horizontal line of the earth is shown by the line X--X and a 
line perpendicular thereto by Y--Y. The line of thrust of the rotor is 
shown by the reference F, the inclination of the thrust line with respect 
to the line Y--Y by the reference .delta., and the inclination of the 
vertical axis through the center of gravity by the reference .theta.. 
In case of a rotor having rotor blade which are movable about flapping 
hinges, each of the blades is subjected to the gravitational force due to 
its own weight, the aeronautical force, the centrifugal force and the 
inertia force and these forces are balanced so that the moment about the 
flapping hinge becomes zero. 
Thus, the following equation can be established. 
EQU .beta.+2T.beta.+.OMEGA..sup.2 .beta.=.OMEGA..sup.2 
.beta.o+2T.OMEGA.(.theta.+.phi.) .sub.sin .psi. (1) 
where .psi. is the angle of rotation of the rotor blade, and .beta. the 
angle of flapping of the blade. 
.OMEGA.: angular velocity of the rotor 
.theta.: the tilting angle of the fuselage; 
From the equation (1), the following approximate equation (2) can be 
obtained provided that the value T/(2.OMEGA.) is larger than zero and 
small in relation to one, and the value (.theta.+.phi.)/.OMEGA.2 is small 
in relation to value (.theta.+.phi.). 
EQU .delta.=T(.theta.+.phi.-.delta.) (2) 
where .delta. is the angle of inclination of the swash plate with respect 
to the rotor mast which is proportional to the angle of actuation of the 
control stick. 
Further, in the system shown in FIG. 3, the following relations are 
established. 
EQU .theta.=K(.delta.-.theta.) (3) 
From the equations (2) and (3); 
EQU .theta.+T.theta.+K.theta.=KT.psi. (4) 
Considering a case wherein the control stick or the swash plate is actuated 
from zero to an angle .psi.o, the solution of the equation (4) becomes as 
follows: where T.sup.2 /4 is small in relation to K, that is, T.sup.2 
/4&lt;&lt;K, 
##EQU3## 
where T.sup.2 /4 is large in relation to K, that is, T.sup.2 /4&gt;&gt;K, 
##EQU4## 
In the equations (5), since the value T/2 is considered as being 
relatively large, the angular velocity .theta. reaches a constant value 
T.psi.o within a relatively short period. As one example of the present 
invention, the value T may be 4 and the value K may be 50. In this case, 
the angular velocity .theta. changes as shown by the curve (a) in FIG. 5. 
However, in the equations (6), since the value K/T is considered as being 
very small and nearly zero, the angular velocity .theta. can reach the 
aforementioned constant value T.psi.o only after a substantial time. In a 
typical example of a conventional helicopter, the value T may be 20 and 
the value K may be 4. In this case, the angular velocity .theta. changes 
as shown by the curve (b) in FIG. 5. 
It will therefore be seen that, in the case where the value T.sup.2 /4 is 
small in relation to K, the ultimate angular velocity of the movement of 
the helicopter fuselage as obtained in response to the actuation of the 
swash plate by a predetermined angle is smaller than that in the case 
where the value T.sup.2 /4 is large in relation to the value K, and the 
ultimate angular velocity can be attained much faster in the former case 
than in the latter case. 
Referring now to the inclination of the rotor with respect to the rotor 
mast, which may be represented by the value (.delta.-.theta.), the 
following relations can be obtained from the equations (3), (5) and (6), 
where the value T.sup.2 /4 is small in relation to the value K, that is, 
T.sup.2 /4&lt;&lt;K, 
##EQU5## 
where the value T.sup.2 /4 is large in relation to the value K, that is, 
T.sup.2 /4&gt;&gt;K, 
##EQU6## 
In the equation (7), the value T/2 is considered as being relatively 
large, the value (.delta.-.theta.)/.phi.o is relatively rapidly damped to 
zero. This means that, with the condition under discussion, the angle of 
inclination of the rotor with respect to the rotor mast can be maintained 
very small. 
In the equation (8), however, the value K/T is considered as being very 
small and nearly zero, the value (.delta.-.theta.)/.phi.o decreases to 
zero only after a substantial time. This will mean that when the swash 
plate is actuated through the control stick by the predetermined angle 
.phi.o, the rotor plane is at first tilted with respect to the rotor mast 
and, only after a substantial time delay, the fuselage with the rotor mast 
starts to move in the direction of the rotor inclination. 
Thus, in the former case, that is, in one example of the present invention, 
the fuselage of the helicopter quickly responds to the actuation of the 
swash plate. The movement .delta. of the rotor plane and that .theta. of 
the fuselage are shown in FIG. 4. In the drawing, the dotted line C shows 
the angle .delta. in accordance with said one example of the present 
invention, the solid line d the angle .theta. of the example, the dotted 
line e the angle .delta. in the conventional one and the solid line f the 
angle .theta. in the conventional one. 
In FIG. 4, it will be seen that, in the example of the present invention, 
the fuselage of the helicopter quickly responds to the inclining movement 
of the rotor plane. Thus, the controllability of the helicopter is 
appreciably improved. Although the discussion has been made with respect 
to a specific example wherein the value T.sup.2 /4 is much smaller than 
the value K, the inventor has found that a significant improvement on the 
controllability can be accomplished with the ratio 2.sqroot.K/T greater 
than 0.8, more preferably, greater than 1.0. 
Referring now to FIGS. 6 through 9 of the drawings which show one practical 
embodiment of the present invention, the helicopter shown therein includes 
a fuselage 11 constituted by a pipe-framework and having an engine 12 
mounted thereon. The fuselage 11 further has a hollow mast 13 which is 
secured thereto so as to extend substantially vertically. A rotor drive 
shaft 23 extends coaxially in the mast 13 and operatively connected with 
the engine 12 through an appropriate clutch and transmission mechanism 
(not shown) so as to be driven by the engine. 
A tail boom 14 extends from the fuselage toward aft and supports a tail 
rotor 15. The tail rotor 15 is connected through a drive shaft 25 with the 
engine 12 to be driven thereby. 
At the upper end of the rotor drive shaft 23, there is mounted a rotor hub 
16 which carries a pair of diametrically opposed rotor blades 17. 
Referring more specifically to FIG. 7, the rotor drive shaft 23 has a 
further shaft 33 which is secured to the upper end thereof so as to extend 
upwardly. At the upper end of the shaft 33, there is secured the 
aforementioned rotor hub 16 of substantially rectangular shape. Each rotor 
blade 17 is connected to the rotor hub 16 through a connecting member 40. 
As shown in FIG. 9, the connecting member 40 is connected through a 
flapping hinge pin 41 to one end of the rotor hub 16. The pin 41 extends 
substantially in a horizontal plane perpendicularly to the longitudinal 
axis of the rotor blade 17 so that the connecting member 40 is movable or 
swingable about the pin 41 in a substantially vertical plane. 
The rotor blade 17 is provided at the end adjacent to the rotor hub 16 with 
a stub shaft 42 which extends radially inwardly of the rotor and is 
received by a bearing 43 for feathering movement about its longitudinal 
axis. Further, the stub shaft 42 is maintained against axial movement by 
means of a thrust bearing 44 and a nut 45. Thus, it will be seen that the 
rotor blade 17 is capable of feathering movement about its longitudinal 
axis and of flapping movement about the hinge pin 41. It should further be 
noted that the flapping hinge 41 is offset from the shaft 33 by a 
substantial distance. 
On the shaft 33, there is mounted a sleeve 46 which is maintained 
stationary as hereinafter be described. For the purpose, the sleeve 46 is 
mounted on the shaft 33 through bearings 47 and 48. The sleeve 46 carries 
a part-spherical bearing 49 on which a swash plate 50 is mounted for 
inclining movement with respect thereto. The swash plate 50 carries a 
follower ring 51 which is rotatable by being mounted on the swash plate 50 
through bearings 52, 53 and 54. The ring 51 has a pair of diametrically 
opposed lugs 55, each of which is connected through a feathering rod 56 
with an arm 57 provided on the corresponding rotor blade 17. 
Referring back to FIG. 6, the helicopter is provided with a collective 
pitch control lever 58 which is connected through a bell-crank 59 and a 
push-pull rod 60 with the sleeve 46 in such a manner that the sleeve 46 is 
vertically displaced through an actuation of the lever 58. Thus, it will 
be understood that, by effecting the vertical displacement of the sleeve 
46 through the actuation of the control lever 58, the swash plate 50 and 
the ring 51 are also shifted vertically so that the leading edges of the 
rotor blades 17 are also vertically moved to change the pitch angles 
thereof. 
The helicopter is further provided with a cyclic pitch control lever 61 
which is mounted at one end on the sleeve 46 for lateral and longitudinal 
swinging movements. More specifically, the sleeve 46 is formed with a lug 
62 which is connected through a spherical bearing 63 and a bolt 64 
extending therethrough with the upper end of the lever 61. The upper end 
of the lever 61 has a rearward extension 65 which is connected through a 
rod 66 with the swash plate 50 at a point 67 located at the left side 
portion thereof. The upper end of the lever 61 is further connected at a 
point beneath the bolt 64 with a connecting rod 68 which is in turn 
connected through a bell-crank 69 and a further rod 70 with the swash 
plate 50 at a point 71 which is circumferentially spaced from the point 67 
by an angle of 90.degree.. 
Thus, it will be understood that by laterally swinging the lever 61, the 
swash plate 50 is longitudinally inclined causing lateral tilting of the 
rotor plane as well known in the art. When the lever 61 is swung 
longitudinally, the swash plate 50 is inclined laterally so that a 
longitudinal tilting movement is produced in the rotor plane. 
The lever 61 is further provided with a throttle control grip 72 which is 
adapted to actuate the throttle valve (not shown) of the engine 12 through 
a cable assembly 73. Further, the lever 61 is provided with a clutch 
actuating lever 74 for controlling the clutch between the engine 12 and 
the drive shaft 23. 
In the embodiment, the rotor blades 17 may be constructed as shown in FIG. 
2 from aluminum alloy by extrusion technique. Alternatively, they may be 
manufactured through any known method provided that the aforementioned 
requirements are met. 
By way of examples, practical values in accordance with the present 
invention are compared with those in conventional helicopters in the 
following table. 
As previously described, in accordance with the present invention, an 
improved controllability can be obtained by maintaining the ratio 
2.sqroot.K/T greater than 0.8, and a further improvement can be 
accomplished with the ratio 2.sqroot.K/T greater than one. According to 
the analysis made by the inventor, when the ratio 2.sqroot.K/T is 
determined as being greater than one, the tilt angle between the rotor 
mast and the axis of the rotor plane quickly converges to zero with 
certain hunting oscillations, so that the fuselage responds to the 
actuation of the swash plate very quickly. 
__________________________________________________________________________ 
##STR1## 
##STR2## 
##STR3## 
(m)(mm)RC 
(1) (rpm)b.OMEGA. 
(kg)W 
(kg.multidot.m.multidot.s.sup.2)I 
(m) (mm)hH 
(10.sup.3 kg)Zo 
(kg.multidot.m.multidot.s.sup. 
2) (m)I.sub.R e 
##STR4## 
##STR5## 
__________________________________________________________________________ 
Prior Art 
I 4 20 0.20 5.35 290 
2322 930 
288 1.37 18.50 0 1274 
II 
8 15 0.38 5.08 330 
2382 1450 
307 1.73 25.70 0 2509 
Invention 
I 50 4 3.54 3.5 140 
3293 270 
15.8 1.2 2013 
2.41 
7.210.13 
470 3243 
II 
30 10 1.095 
4 140 
2380 250 
21.9 1.2 3.58 
6.340.1 
358 3000.8 
__________________________________________________________________________ 
The invention has thus been shown and described with reference to specific 
examples, however, it should be noted that the invention is in no way 
limited to the details of the illustrated structures but changes and 
modifications may be made without departing from the scope of the appended 
claims. For example, although the foregoing descriptions have been made 
with respect to rotor systems having two rotor blades, the invention can 
also be applied to rotary wing aircrafts having rotor systems with three 
or more blades. Further, the concepts of the invention can also be applied 
to rotary wing aircrafts having rotor systems in which rotor blades are 
mounted for dragging movements.