Helicopter rotor speed changing transmission

A helicopter has two engines adapted to drive a main sustaining rotor and a plurality of auxiliary propulsion means through a transmission system that is selectively operable to vary the relative speeds of rotation of the sustaining rotor and the auxiliary propulsion means depending upon operating conditions.

This invention relates to helicopters. 
A limiting factor in increasing the forward speed of conventional 
helicopters, in which a main sustaining rotor remains substantially 
edgewise to the airflow during forward flight, is the requirement to 
prevent the advancing blade exceeding its critical mach number, and it is 
recognised that in order to combine this requirement with an efficient 
hover performance it is desirable to be able to vary the rotational speed 
of the rotor. Thus, a high rotational speed is required in hover and a 
lower rotational speed is required in cruise. 
A moderate range of rotor rotational speeds can be obtained by allowing the 
power plant output (usually a free turbine) to run within a range of 
speeds spanning the optimum design speed. However, the range of obtainable 
rotor rotational speeds is limited by the need to ensure that the 
inevitable loss of engine efficiency is not excessive. 
It is also recognised in the art that a high speed helicopter will require 
an auxiliary propulsion means such as a propeller or fan in order to 
supplement the propulsive capability of the main sustaining rotor which 
reduces in efficiency with increasing forward speed. In such an 
arrangement it has been found that the distribution of power between the 
main sustaining rotor and the auxiliary propulsion means varies greatly 
between the requirements of hover and cruise flight, in that in hover the 
power is required to be directed mainly to the sustaining rotor whereas in 
cruise it is required to be directed mainly to the auxiliary propulsion 
means. It follows, therefore, that in installations in which separate 
engines are provided for driving the sustaining rotor and the propulsion 
means, the total power capability will be under-employed throughout the 
speed range of the helicopter. With gas turbines this will result in high 
fuel consumption from a large and expensive power plant installation. 
An alternative arrangement is to shaft drive both the sustaining rotor and 
the auxiliary propulsion means from the same power plant. This provides 
for lower fuel consumption from less installed power but, however, 
problems then arise from the need to reduce the rotational speed of the 
sustaining rotor with increasing cruise speed. In a simple transmission 
system it follows that such a speed reduction of the sustaining rotor will 
be accompanied by a reduction in the speed of the auxiliary propulsion 
means at the very time that the speed of the latter needs to be 
increasing. If the gear ratio between the power plant and the auxiliary 
propulsion means is selected to provide for the higher speed more suited 
to cruise flight then the auxiliary propulsion means will overspeed during 
hover, creating excessive noise and using excessive power. 
A prior proposal concerned with these problems is disclosed in U.S. patent 
application No. 2,665,859. In that proposal an aircraft is provided with 
both fixed and rotary wings and a pair of laterally spaced propellers 
located on engine nacelles on the fixed wings. A transmission unit is 
associated with each of the engines and includes a brake band clutch 
device for selectively controlling the transmission of power to the 
propeller and two brake band clutch devices associated with two different 
gear ratios for selectively transmitting power at a selected ratio to the 
rotary wings. 
It is to be noted that in this prior arrangement when the respective brake 
band clutches are disengaged no power is transmitted so that for instance, 
during hover, the propellers are stationary. Consequently when it is 
desired to engage the propellers it is firstly necessary to actuate both 
brake band clutches simultaneously and secondly, since it is impossible to 
translate instantly from zero to full speed, a great deal of sliding of 
the clutch parts will occur resulting in the generation of tremendous heat 
and consequent heat sink problems. This problem is exacerbated by the fact 
that there are six brake band clutches in the installation, giving rise to 
problems in manipulating large numbers of such devices and in achieving 
required synchronisation of the operation of pairs or groups of these 
devices. 
The requirement to manipulate four brake band clutches during attempts to 
change the rotor speed could also result in an inability to control the 
rotor speed whilst changing from one gear ratio to the other. 
Accordingly, in one aspect, a helicopter having two power sources for 
driving a main sustaining rotor and an auxiliary propulsion means includes 
a transmission system selectively operable to vary the relative speeds of 
rotation of the sustaining rotor and the auxiliary propulsion means, the 
transmission system including, for each power source, a layshaft, a first 
output pinion on each said layshaft connected through a freewheel device 
to a rotor output gear for driving said main sustaining rotor, a second 
output pinion on each said layshaft connected through a freewheel device 
to an auxiliary propulsion output gear for driving said auxiliary 
propulsion means, and a third output pinion on each said layshaft 
drivingly connected to a common differential gear having a central output 
shaft connected through an actuatable freewheel clutch to said rotor 
output gear. 
In a modified form of the invention the central output shaft may be 
connected through an actuatable freewheel clutch to the auxiliary 
propulsion output gear, and in a further modified form the central output 
shaft may be connected through actuatable freewheel clutches to both the 
rotor output gear and the auxiliary propulsion output gear. 
Preferably the gear ratio between said third output pinion and said 
differential gear is lower than that between said first and second output 
pinions and the respective rotor and auxiliary propulsion output gears. 
The main sustaining rotor may comprise a co-axial rotor and may be driven 
from said rotor output gear through a planetary gear system. 
The auxiliary propulsion means may comprise two propellers driven from said 
auxiliary propulsion output gear through suitable gearing and drive 
shafts. 
In another aspect the invention provides a helicopter having two engines 
arranged to drive a main sustaining rotor and auxiliary propulsion means 
through a transmission system wherein the transmission system includes 
layshafts driven by the respective engines, each layshaft having a first 
output gear drivingly connected through a freewheel device to a rotor 
output gear for driving the sustaining rotor, a second output pinion 
drivingly connected through a freewheel device to an auxiliary propulsion 
means output gear for driving said auxiliary propulsion means, and a third 
output pinion drivingly connected to a common in-line differential gear 
having a central output shaft connected through actuatable freewheel 
clutches to the rotor output gear and the auxiliary propulsion means 
output gear, the gear ratio between the third output pinions and the 
differential gear being lower than that between the first and second 
output pinions and the rotor output gear and auxiliary propulsion means 
output gear respectively.

A helicopter 39 (FIG. 1) includes two power sources comprising port and 
starboard engines 40 (one only being shown) and a transmission system 
(FIGS. 2 to 5 inclusive) for connecting the engines to drive a main 
sustaining rotor 41, comprising two contra-rotating co-axial rotors, and 
two auxiliary propellers 42 (one only being shown) located respectively at 
the outer ends of stub wings 43 extending outwardly from a fuselage 44. 
The transmission system generally indicated at 11 in FIG. 2 includes port 
and starboard engine input shafts 12 and 13 connected through bevel gear 
sets 14 and 15 to port and starboard layshafts 16 and 17 respectively. 
Each layshaft 16 and 17 has a first (upper) output pinion 16a and 17a 
connected through freewheel clutches 18 and 19, a second (lower) output 
pinion 16c and 17c connected through freewheel clutches 20 and 21 and a 
fixed third (central) output pinion 16b and 17b. Freewheel clutches 18, 
19, 20 and 21 are adapted to transmit torque from the layshafts 16 and 17 
to respective output gears, and to overrun when the speed of rotation of 
the output gears exceeds that of the respective layshafts 16 and 17. 
The first pinions 16a and 17a engage a rotor output gear 22 which drives 
the sun gear 23 of a planetary gear system generally indicated at 24 
adapted to drive the co-axial drive shafts 25 and 26 connected to the main 
sustaining rotor 41 (FIG. 1). The second pinions 16c and 17c engage an 
auxiliary propulsor output gear 27 which is connected through bevel gear 
sets 28, shafts 29 and 30, and further bevel gear sets 31 and 32 to drive 
port and starboard propulsor drive shafts 33 and 34 respectively connected 
to the twin auxiliary propellers 42 (FIG. 1). 
The third pinions 16b and 17b are staggered axially on the layshafts 16 and 
17 to enable them to engage a common centrally positioned in-line 
differential gear unit 35. Differential gear unit 35 is mounted on a 
central shaft 36 which provides during certain phases of operation a 
second drive path to output gear 22 through a lockout or actuatable 
freewheel clutch 37, and a second drive path to output gear 27 through a 
further lockout or actuatable freewheel clutch 38. 
Actuatable freewheel clutches 37 and 38 are selectively operable between 
position `I` in which torque can be transmitted from central shaft 36 to 
the respective output gears 22 and 27 and position `O` in which no torque 
can be transmitted, and it will be understood that clutches 37 and 38 will 
over-run even in engaged position `I` should the speed of rotation of the 
respective output gears 22 and 27 exceed that of central shaft 36. The 
mode of operation of actuatable clutches 37 and 38 cannot be changed while 
torque is being transmitted and the main purpose of the in-line 
differential gear unit 35 is to unload the clutches 37 and 38 during 
transition phases of operation as described in detail hereinafter. 
It will be noted that the third pinions 16b and 17b are larger in diameter 
than the respective first and second pinions 16a, 17a, 16c and 17c to 
provide a correspondingly lower gear ratio and therefore a higher output 
drive speed when connected either to the rotor output gear 22, or the 
auxiliary propulsor output gear 27. 
Operation of a helicopter incorporating the transmission system of this 
invention will now be described with reference to FIGS. 3, 4 and 5, in 
which like reference numerals have been used to indicate parts identical 
to those previously described. In all cases the high speed drive path is 
indicated by the uniform broken line and the low speed drive path by the 
chain dot line. 
FIG. 3 represents the transmission configuration for take off, landing, 
hover and low speed forward flight. In this configuration actuatable 
clutch 38 is at position `O` and actuatable clutch 37 is at position `I`. 
Torque from input shafts 12 and 13, is transmitted to port and starboard 
layshafts 16 and 17. Due to the lower gear ratio between third output 
pinions 16b and 17b, and the inline differential gear unit 35, the high 
speed drive is transmitted through central shaft 36, actuatable clutch 37, 
output gear 22 and planetary gear 24, to drive co-axial drive shafts 25 
and 26 attached to the main sustaining rotor 41 (FIG. 1). 
Due to the higher gear ratios, freewheel clutches 18 and 19 overrun during 
this phase of operation; however, since actuatable clutch 38 is at 
position `O`, low speed drive is transmitted through freewheel clutches 20 
and 21 to output gear 27, to drive the drive shafts 33 and 34 of the 
auxiliary propellers 42 (FIG. 1). 
Thus, during this phase of operation the main sustaining rotor 41 is driven 
at high speed to obtain adequate vertical thrust, while the auxiliary 
propellers 42 are driven at low speed to conserve power consumption. 
When it is desired to change from low to high forward speed and vice versa 
it is necessary to enter a transition phase to enable the mode of 
operation of actuatable clutches 37 and 38 to be changed. 
Considering firstly transition from the hover/low speed operation described 
with reference to FIG. 3 to a high forward speed operation it is necessary 
firstly to reduce rotor r.p.m. by re-setting the helicopter rotor speed 
governing datum. The power from one engine, say the starboard engine 
connected to input shaft 13, is then reduced to a flight idle setting, and 
the consequent reduction in speed of the starboard layshaft 17 will 
automatically disengage it immediately from output gear 27 by overrunning 
of clutch 21 associated with the second pinion 17c. 
This initial reduction in the speed of rotation of starboard layshaft 17 
creates a difference in speed between the two input gears 16b, 17b of the 
in-line differential gear 35. The central shaft 36 of the differential 
gear 35 is therefore rotated at the mean speed of the two inputs and will 
reduce correspondingly the speed of rotation of the output gear 22, and 
therefore the main rotor drive shafts 25 and 26. 
The reduction in main rotor speed will be sensed by the helicopter rotor 
speed governor which will automatically increase the power output from the 
port engine to maintain datum rotor speed, and this will increase the 
speed of rotation of port layshaft 16 until it catches up with the speed 
of rotation of first pinion 16a to engage through clutch 18 with output 
gear 22. Further increase in speed of layshaft 16 increases the speed 
differential with layshaft 17 causing actuatable clutch 37 to disengage by 
overrunning. Clutches 19 and 21 will overrun leaving the drive path to the 
main sustaining rotor 41 entirely through clutch 18 and the drive path to 
the auxiliary propellers 42 through clutch 20 as indicated by the dotted 
lines in FIG. 4. 
Thus, at this interim stage of the transition operation, a drive path to 
the main sustaining rotor and its speed have been maintained, the speed of 
operation of the port engine and the drive shafts 33 and 34 of the 
auxiliary propellers 42 has been increased, and both of the actuatable 
clutches 37 and 38 are automatically unloaded and disengaged as indicated 
by the intermediate positions indicated in FIG. 4. 
Actuatable clutch 38 is now moved to engage mode `I` and actuatable clutch 
37 to lockout mode `O` (FIG. 5). Power from the starboard engine is now 
increased to match that of the port engine, and the corresponding increase 
in the speed of rotation of starboard layshaft 17 increases the speed of 
rotation of the central shaft 36 of differential gear 35 until it catches 
up with the speed of rotation of output gear 27 which will engage 
actuatable clutch 38. A further increase in the speed of rotation of the 
starboard layshaft 17 will further increase the speed of rotation of 
output gear 27 and, correspondingly, the auxiliary propellers 42. Due to 
the different gear ratios, clutches 20 and 21 will overrun leaving the 
drive path to the propellers 42 entirely through the low gear ratio of 
third pinions 16b and 17b and actuatable clutch 38, as indicated by the 
high speed drive path shown in uniform broken line in FIG. 5. 
As the speed of rotation of the starboard layshaft 17 reaches that of the 
port layshaft 16, clutch 19 will engage to enable the starboard engine to 
contribute its power to drive the main rotor drive shafts 25 and 26 
through the low speed drive path indicated by the chain dot line in FIG. 
5. 
This completes the transition to leave the main sustaining rotor 41 at a 
low speed, and the auxiliary propellers 42 at a high speed, suitable for 
high speed forward flight. It will be noted that following this initial 
transition mode of the described embodiment, the engine free turbines are 
running at a high speed and such an arrangement permits a further 
reduction of main rotor speed as forward flying speed increases by 
progressively resetting the helicopter engine governor datum to reduce 
free turbine speeds. 
Transition from high to low forward speed drive ratios is the reverse of 
the above. Briefly, as forward speed reduces towards minimum power speed, 
the main sustaining rotor speed is increased by increasing the engine free 
turbine speed. The power of one of the engines is then reduced to a flight 
idle setting which in turn will reduce the speed of rotation of the 
auxiliary propulsor output gear 27 and unload actuatable clutch 38. 
Actuatable clutch 38 is then changed to lockout mode (position `O`) and 
actuatable clutch 37 is changed to engage mode (position `I`), and engine 
power is restored. This engages clutch 37 to revert to the respective 
drive paths illustrated in FIG. 3. 
Selectively reducing the rotor speed governor decreases the engine free 
turbine speeds to further reduce the speed of auxiliary propulsor output 
gear 27, and this arrangement permits a further increase in main 
sustaining rotor speed in the hover mode by resetting the engine governor 
datum to increase free turbine speeds. 
For descriptive purposes the transition procedures have been explained in 
detail and may appear rather complex; however, in practice it is to be 
noted that it is only necessary for the pilot to reduce the power from one 
or other of the engines, select the correct mode of operation of the 
respective actuatable freewheel clutches 37 and 38 and restore the power 
from the engine. 
In contrast with the prior proposal of U.S. application Ser. No. 2,665,859 
the present invention utilises a single transmission unit for controlling 
the relative speeds of a main sustaining rotor and auxiliary propulsion 
means from power derived from two engines rather than a transmission unit 
associated with each engine. This results in a much lighter, simpler 
system since it requires operation of two devices only to effect speed 
changes compared to the six devices of the prior proposal. Furthermore, 
operationally, the use of actuatable freewheel type clutches in the 
present invention is advantageous firstly since engagement occurs 
automatically at synchronous speeds with none of the tremendous heat 
generation of the prior proposal and, secondly, since transition can be 
accomplished without the requirement to physically disconnect the drive, 
control over the relative speeds of the sustaining rotor and the auxiliary 
propulsion means is retained at all times. 
Whilst one embodiment of the invention has been described and illustrated 
it will be apparent that many modifications may be made without departing 
from the scope of the invention as defined in the appended claims. For 
example, in installations requiring a smaller range of available operating 
speeds between the main sustaining rotor and the auxiliary propulsion 
means one or other of the lockout freewheel clutches can be omitted so 
that major speed variation is accomplished individually in either the main 
sustaining rotor or the auxiliary propulsion means. The invention can be 
incorporated in a helicopter having a single main sustaining rotor as 
opposed to the co-axial rotor of the illustrated embodiment although some 
form of anti-torque means would then be required. Alternatively, the 
anti-torque force could also be provided by one or both of the propellers 
42 due to their lateral location on auxiliary wing structures 43. The 
auxiliary propulsion means may comprise propulsive fans, and more or less 
than the two devices of the illustrated embodiment can be provided. Each 
of the two power sources may comprise more than one engine, the output 
from the engines of each power source either being combined to drive the 
respective single input shafts 12 and 13 or driving individual input 
shafts for engagement respectively with the two layshafts 16 and 17.