Helicopters

A helicopter including a power absorber arranged to selectively absorb power from a sustaining rotor system during deceleration. In an illustrated embodiment for a helicopter having a single sustaining rotor, the power absorber comprises a tail rotor assembly having two tail rotors located one on each side of a tail cone, and a control operative to adjust the pitch setting of the tail rotors to provide lateral thrust forces in opposed directions with a net lateral thrust force in one direction in order to maintain a desired heading during deceleration.

DESCRIPTION OF INVENTION 
This invention relates to helicopters. 
High forward speeds attainable by modern helicopters have resulted in 
corresponding increases in the required stopping distances, especially in 
respect of straight line stops from high speed level flight. 
The procedure in such a manoeuvre is for the pilot to reduce power and 
operate the cyclic control to rotate the helicopter into a nose-up 
attitude. The forward velocity of the helicopter creates a reverse air 
flow through the rotor(s) of the sustaining rotor system which tends to 
increase the rotational speed of the rotor and this is usually controlled 
by a rotor speed governing system acting automatically to reduce the 
engine power output to the rotor to maintain a constant rotor speed. As 
the nose-up attitude is further increased the engine power output 
continues to be decreased until, ultimately, the engine power being 
supplied to the rotor system approaches zero. This sets a limit on the 
attainable nose-up attitude since a further increase would increase the 
reverse airflow causing an increase in rotor speed beyond an acceptable 
limit. 
Thus, helicopter straight-line deceleration from high speed forward flight 
is limited by the nose-up attitude to which the helicopter can be rotated. 
In the case of a helicopter having a single main sustaining rotor and an 
anti-torque auxiliary rotor, e.g. a tail rotor, the limitation on nose-up 
attitude occurs when the main rotor tends to overspeed as a result of 
"autorotation" of the rotor by the reverse mass airflow. High nose-up 
attitudes can be attained only at low forward speeds and do not 
significantly reduce an overall stopping distance. At high forward speeds 
the fuselage parasitic drag may allow adequate initial deceleration but 
this becomes progressively less effective as the helicopter slows down. 
Accordingly, in one aspect, the invention provides a helicopter 
characterised by power absorber means arranged to selectively absorb power 
from a sustaining rotor system during deceleration. 
The power absorber means may comprise at least one pair of auxiliary 
rotors, and the helicopter may have a control system arranged during 
normal operation to adjust the pitch setting of each pair of auxiliary 
rotors to provide thrust forces in the same direction and, during 
deceleration, to adjust the pitch setting of each pair of auxiliary rotors 
to provide thrust forces in generally opposed directions. 
Although the invention is applicable to helicopters of other 
configurations, e.g. those with main sustaining rotor systems comprising 
two or more rotors, such as tandem and co-axial rotor systems, it is 
especially advantageous in helicopters having a single main sustaining 
rotor and an anti-torque auxiliary rotor, e.g., a tail rotor. In applying 
the invention to a helicopter of this configuration, a pair of auxiliary 
rotors may be located adjacent the end of a rearwardly extending tail 
cone. Conveniently, the auxiliary rotors are supported laterally one on 
each side of the tail cone and are arranged so that said thrust forces 
comprise lateral thrust forces in a yaw direction so as to operate as an 
anti-torque tail rotor during normal operation. The direction of the 
opposed thrust forces may be generally inwardly from each auxiliary rotor. 
The helicopter may include a differential pitch adjusting mechanism 
activated automatically at the onset of zero torque to the single main 
sustaining rotor by electrical signals from a rotor speed governor and a 
yaw/heading control mechanism to adjust the pitch setting of the auxiliary 
rotors differentially to provide a net lateral thrust force in a desired 
yaw direction in order to maintain the helicopter on a desired heading. 
The yaw/heading control mechanism may be controlled by rudder pedals 
incorporating a gearing mechanism operative to vary the yaw demand 
dependant on the actual collective pitch setting of the tail rotors, the 
gearing being phased-in automatically by an electrical signal generated by 
the yaw/heading control mechanism in response to a reference signal from 
the differential pitch adjusting mechanism. 
Preferably, a manual inhibition system is arranged to selectively isolate 
the differential pitch adjusting mechanism so as to prevent differential 
pitch being applied during certain phases of operation. 
In another aspect the invention provides a helicopter having a fuselage 
including a rearwardly extending tail cone, a main sustaining rotor 
located above the fuselage for rotation about a generally vertical axis, a 
tail rotor assembly located adjacent a rear end of the tail cone and 
comprising two tail rotors supported laterally one on each side of the 
tail cone, and a control system arranged during deceleration of the 
helicopter to adjust the pitch setting of both tail rotors so that the 
tail rotors provide respective lateral thrust forces in generally opposed 
directions. 
Preferably, the control system includes a differential pitch adjusting 
mechanism arranged to control the pitch setting of the two tail rotors 
differentially to provide unequal lateral thrust forces in the generally 
opposed directions resulting in a variable net lateral thrust in one 
desired lateral direction. 
As previously explained, during deceleration a helicopter having a single 
main sustaining rotor operates in an autorotative condition in which the 
rotor is rotated by reverse airflow through the main rotor disc. 
Autorotation is, of course, also used in helicopter operation to enable a 
helicopter to land safely in the event of a power failure, however, it is 
the autorotative condition existing during the deceleration mode of 
operation with which this invention is particularly concerned.

In considering the problem of improving the deceleration characteristics of 
a high speed helicopter having a single main sustaining rotor, the 
inventors recognised that it would be necessary to increase the attainable 
nose-up attitude without a corresponding increase in the rotor rotational 
speed. In this "autorotative" condition, power to maintain rotation of the 
main rotor is provided by reverse airflow through the rotor disc, and an 
accessory gearbox is normally driven by the main rotor to maintain control 
of the helicopter. The inventors realised that, if the main rotor could be 
made to do more work in this condition, then more power could be absorbed 
by the main rotor before rotor overspeed occurred, and that a greater 
reverse mass airflow could be tolerated thereby increasing the attainable 
nose-up attitude and, consequently, the deceleration force being applied 
by the main rotor. 
In normal flight, torque applied to a main rotor of a helicopter by a power 
source creates a turning moment tending to rotate a fuselage about the 
main rotor rotational axis, and this is reacted by a generally lateral 
thrust force from a tail rotor. 
In the autorotative condition the main rotor is driven by reverse airflow 
and, in turn, provides the power to drive an accessory gearbox. This 
results in a relatively small turning moment on the fuselage which is 
reacted by a small residual lateral thrust from the tail rotor. 
Since the deceleration manoeuvre takes the helicopter into an autorotative 
condition the tail rotor is normally feathered during the manoeuvre so as 
to provide only the small residual lateral thrust required to compensate 
for the torque reaction produced by the main rotor. 
At this time, the inventors were also involved in an appraisal of various 
tail rotor configurations with the aim of improving normal yaw control of 
a helicopter, and to provide continued control in the event of tail rotor 
loss, and one of the schemes under consideration was a twin tail rotor 
configuration consisting of two tail rotors mounted on fins or outriggers 
laterally of the tail cone. However, the inventors realised that potential 
in utilising the twin tail rotor concept as the means of absorbing power 
from the main rotor during the autorotative condition experienced in the 
deceleration manoeuvre, by arguing that if the two tail rotors could be 
arranged to provide high thrust in opposite directions whilst retaining 
the required small net lateral thrust, then this should extract a 
significant amount of extra power from the main rotor and allow a higher 
nose-up attitude to be attained before rotor overspeed occurred. 
Subsequent calculations indicated that, for a particular helicopter having 
a single main sustaining rotor, the incorporation of a twin tail rotor 
having the facility to provide opposed thrust could reduce the stopping 
distance of that helicopter by as much as one third. This represents a 
useful benefit in respect of the required stopping distance, and is a very 
attractive solution since it can be combined in a single mechanism which 
provides other advantages such as the improved yaw control and tail rotor 
loss capability of the twin tail rotor configuration. 
Referring now to FIGS. 1 to 3 inclusive, a helicopter has a fuselage 10 
including a rearwardly extending tail cone 11. A single main sustaining 
rotor 17 is mounted above the fuselage 10 and a tail rotor assembly, 
generally indicated at 12, is located adjacent a rear end of the tail cone 
11. The tail rotor assembly 12 comprises two tail rotors 13 located 
respectively at the ends of laterally extending fins 14 and in the 
illustrated embodiment, the fins 14 extend upwardly and outwardly from the 
tail cone 11 at an angle of approximately 20 degrees. Each tail rotor 13 
is mounted for rotation in a plane perpendicular to its respective fin 14. 
One or more engines and a transmission system (not shown) are provided and 
arranged to drive the main sustaining rotor 17 and the two tail rotors 13 
through suitable drive shafts and gearing. 
The blades of each tail rotor 13 are adjustable collectively in pitch to 
provide a lateral thrust in either direction perpendicular to their plane 
of rotation, and the helicopter includes a control system for adjusting 
the pitch of each of the individual tail rotors 13 to provide either 
lateral thrust forces in the same direction, or varying degrees of lateral 
thrust forces in opposite directions. 
A control system for controlling uniform adjustment of the tail rotors 13, 
i.e. during normal flight in which the tail rotors 13 are employed 
conventionally to provide yaw control, does not form part of this 
invention and is therefore not described in detail in the following 
description. However, it will be understood that certain features of the 
control system now to be described with reference to the present invention 
will also have a functional requirement during conventional control 
manoeuvres. Such equipment includes a rotor speed governor, a yaw/heading 
control mechanism and rudder pedals. 
Referring now to FIG. 4, one or more engines 15 are arranged to supply the 
required torque through drive shafts and gearing 16 to drive the main 
rotor 17, and through further shafts and gearing 18 and a tail rotor 
differential pitch adjusting mechanism 19 to drive the tail rotor assembly 
12. 
Electrical signals 25 representative of the actual torque being applied to 
the main rotor 17 are transmitted to a rotor speed governor 20 and a 
failure detection system 21, and a signal 26 generated by the rotor speed 
governor 20 and representative of a torque demand is transmitted to the 
engine 15. The rotor speed governor 20 and failure detection system 21 
also receive signals 27 representative of the rotational speed of the main 
rotor 17. 
During a certain phase of operation to be hereinafter described, the rotor 
speed governor 20 generates a demand signal 28, and this is transmitted to 
the tail rotor differential pitch adjusting mechanism 19 which adjusts the 
pitch setting of the two tail rotors 13 to provide lateral thurst forces 
in opposed directions. Signals 29 representative of the applied pitch 
setting are transmitted from the mechanism 19 to the rotor speed governor 
20, the failure detection system 21 and to a yaw/heading control mechanism 
22. 
The yaw/heading control mechanism 22 receives yaw demand signals 30 from 
pilot rudder pedals 23, and generates a signal 31 based on the applied 
pitch setting of the tail rotor assembly 12 to suitably adjust a gearing 
associated with the rudder pedals 23. A signal 32 representative of a yaw 
rate and acceleration is generated by the yaw heading control system 22 
and is transmitted to the failure detection system 21 and to the 
differential pitch adjusting mechanism 19. 
A manually operable inhibition system 24 is provided and is connected 
electrically to the tail rotor differential pitch mechanism 19 for a 
purpose to be hereinafter described. 
When it is required to stop the described helicopter as quickly as possible 
from high speed forward flight, the pilot reduces power being supplied by 
engine 15 and operates a main rotor cyclic pitch control to rotate the 
helicopter into a nose-up attitude. This tends to increase the rotational 
speed of the main rotor 17 due to reverse airflow through the rotor disc 
and this is sensed by the rotor speed governor 20 which acts to further 
reduce the power output of the engine 15 until the torque being supplied 
to the main rotor 17 approaches zero. This phase of the deceleration 
manoeuvre is conventional and, as previously explained, in helicopters 
having single sustaining and anti-torque rotors and conventional control 
system, the position at which zero torque is being applied to the main 
rotor 17 from engine 15 limits the attainable nose-up attitude and 
therefore the deceleration force which can be applied to the helicopter. 
This, in turn, dictates the required stopping distance of the particular 
helicopter from high speed forward flight. 
It will be understood that the helicopter is now in an autorotative 
condition in which the main rotor is being driven by reverse airflow 
through the rotor disc due to the forward velocity of the helicopter, and 
the tail rotor assembly 12 is being driven from the main rotor 17. 
However, in the described helicopter embodying the invention, the onset of 
zero torque to the main rotor 17 is sensed by the rotor speed governor 20 
which generates electrical signal 28. This signal is transmitted to the 
tail rotor differential pitch adjusting mechanism 19 which adjusts the 
collective pitch setting of the individual tail rotors 13 to provide 
lateral thrust forces in opposite directions generally inwardly from each 
of the tail rotors 13. Yaw demand signal 32 is fed to the mechanism 19 and 
results in a desired differential being applied to the respective pitch 
settings of the tail rotors 13 so that a small net lateral thrust in a 
desired yaw direction is maintained in order to compensate for the 
aforementioned main rotor torque reaction, and maintain the helicopter on 
a desired heading. Thus, the arrangement of the present invention, which 
permits this small net lateral thrust to be achieved with a relatively 
high collective pitch angle setting of each of the individual tail rotors 
13, means that the tail rotor assembly 12 absorbs considerable power from 
the main rotor 17 while the helicopter is in the autorotative condition 
during the deceleration manoeuvre. 
This, in turn, tends to reduce the rotational speed of the main rotor, and 
this is retained at a desired level by an increase in the nose-up attitude 
thereby increasing the reverse mass airflow through the rotor disc. 
The pitch setting of the tail rotors 13 is progressively increased 
differentially to a pre-selected maximum at which the tail rotor assembly 
12 absorbs high power from the main rotor 17 whilst maintaining the 
necessary yaw control thereby establishing the maximum attainable nose-up 
attitude of the helicopter and, correspondingly, the maximum deceleration 
force. 
To terminate the manoeuvre, as the forward speed of the helicopter 
decreases, the pilot reduces the cyclic pitch setting of the main rotor to 
restrict the nose-up attitude. A resultant tendency for the rotational 
speed of the main rotor 17 to decrease is sensed by the rotor speed 
governor 20 which automatically adjusts the signal 28 to the tail rotor 
differential pitch adjusting mechanism 19 to reduce the pitch setting of 
the tail rotors 13. At the same time, a signal 29 representative of the 
actual torque being applied to the tail rotors 13 is transmitted to the 
rotor speed governor 20 and, when this reduces to zero, the governor 20 
transmits a demand signal 26 to the engine 15 to increase the torque to 
the main rotor 17 to maintain the rotational speed of the main rotor 17 at 
the level required to ensure that the helicopter remains airborne and 
controllable in a hovering condition at the end of its deceleration mode, 
i.e., in a normal flight mode. 
The yaw/heading control mechanism 22 ensures that the helicopter remains 
controllable and steerable during the deceleration mode. It is known that, 
as a tail rotor approaches its maximum thrust, increases in pitch are less 
effective in generating thrust for yaw control. In the arrangement of the 
present invention in which, as previously explained, the two tail rotors 
13 are used during deceleration to provide lateral thrust in opposite 
directions, as the maximum pitch limit is approached, yaw control is 
achieved by moving one of the tail rotors 13 away from a stalled condition 
to provide a net lateral thrust to control the helicopter in yaw, i.e., 
once full pitch is applied to the two tail rotors 13, yaw demands, applied 
through signal 32, operate on one only of the tail rotors 13 at a time. 
The mechanism 22 includes logic circuitry to determine which of the two 
rotors 13 is adjusted depending on the yaw demand initiated by the pilot 
at the rudder pedals 23. 
Due to the reduced thrust gradient close to the maximum thrust of the tail 
rotors 13, gearing is associated with the rudder pedals 23 to 
automatically compensate for the actual pitch setting of the two tail 
rotors 13. The gearing has a ratio of at least 4:1 and is phased-in 
automatically by signal 31 from the yaw/heading control mechanism 22 
which, in turn, is initiated by a reference signal from the tail rotor 
differential pitch adjusting mechanism 19. 
In normal flight, loss of one of the two tail rotors 13 can be survived due 
to the high thrust capability of the individual tail rotors 13. However, 
if such a loss occurs during the deceleration mode hereinbefore described, 
a large yaw input may occur accompanied by a tendency for the main rotor 
17 to overspeed. In such an event, the failure detection system 21 
operates to initiate remedial action to ensure the safety of the 
helicopter. The system 21 monitors engine torque and main rotor r.p.m. and 
receives input signals 32 and 29 representative of yaw rate and 
acceleration and an actual pitch setting of the tail rotor assembly 12. 
The manual inhibition system 24 serves to isolate the tail rotor 
differential pitch adjusting mechanism 19 through signal 33 and is used 
during normal autorotative descents, i.e. in the event of a complete power 
failure to prevent opposed thrust forces being applied by the tail rotors 
13 due to the reduction of the torque being applied to the main rotor 17 
to zero. This is a necessary safety factor, since, in an autorotative 
descent, all the available power created by reverse airflow rotating the 
main rotor 17 as the helicopter descends is required to control the rate 
of descent and ensure that the inertia in the rotor 17 is sufficient to 
arrest the rate of descent as the helicopter approaches the ground in 
order to effect a safe landing. 
The effects of the automatic power absorption of the twin tail rotor 
assembly 12 of this invention are illustrated in FIGS. 5 to 7 inclusive 
which are derived from calculations based on a single main rotor 
helicopter of about 25,000 pounds all-up-weight, and assuming that the 
power absorbed (P.sub.out) by the two tail rotors 13 is constant at 1,000 
H.P. Each of the graphs include two curves based on a similar helicopter, 
datum curve A representing results based on the helicopter with a 
conventional single tail rotor and control system, and curve B 
representing results based on a helicopter constructed in accordance with 
the described embodiment of the invention. 
FIG. 5 plots the attainable autorotative attitude (.gamma. radians) against 
forward level flight speed. It will be noted that the attainable attitude 
increases rapidly at speeds below about 120 ft./sec. and this is due to 
the inability of the rotor inflow component to overcome the power losses 
of the rotor while maintaining the helicopter in level flight. The 
significant feature as far as the present invention is concerned is the 
appreciable increase in autorotative attitudes attainable above this 
speed. 
FIG. 6 plots deceleration (g) against forward speed and includes a fuselage 
parasitic drag term. Again the significant feature is the increase in 
deceleration indicated by curve B attainable by the helicopter of this 
invention. 
FIG. 7 presents the stopping distances for the two helicopters being 
considered and is derived from FIG. 6. A stopping distance (ft.) is 
plotted against forward speed, and it is assumed that once the helicopter 
has been rotated nose-up so that the attitude of the rotor disc reaches 30 
degrees, power is resumed to the main rotor to maintain this attitude. It 
will be seen that at low forward flight speed, the incorporation of the 
invention has little effect on the total stopping distance and this is due 
to the fact that higher nose-up attitudes are conventionally attainable at 
low forward speeds as previously explained. However, if we consider the 
case at a high forward speed of 250 ft./sec. it will be seen that the 
helicopter without the invention (curve A) requires a stopping distance of 
about 4,020 ft., whereas the helicopter incorporating the invention (curve 
B) requires a stopping distance of about 2,688 ft., i.e. a reduction in 
stopping distance of about 33 percent. This is achieved according to this 
invention by absorbing power from the main rotor 17 during the 
deceleration manoeuvre. The use of a twin tail rotor as the power absorber 
provides the added advantages of improved yaw response during normal 
flight and improved tail rotor loss survivability. 
However, whilst the invention has been particularly described and 
illustrated in respect of a helicopter having a single main sustaining 
rotor and an anti-torque tail rotor, it is to be understood that the 
incorporation of a selectively operable power absorber means to absorb 
power from a sustaining rotor system as taught by this invention will 
reduce the stopping distance of other configurations such as helicopters 
having tandem or co-axial sustaining rotors, by permitting a higher 
nose-up attitude to be attained during deceleration. Since anti-torque 
rotors are not normally required in such configurations, one or more pairs 
of rotors constituting a power absorber means as in the described 
embodiment may be utilised as auxiliary rotors, and may be arranged to 
provide other functions such as lift, control or propulsion during normal 
operation whilst maintaining the facility during deceleration to provide 
thrust forces in generally opposed directions thereby combining high power 
absorption with no net thrust input. 
The power absorption characteristics of the two rotors 13 of the described 
embodiment can be enhanced by detail design features of the rotor blades 
employed. Thus, an increased blade chord dimension would increase both 
rotor thrust and profile drag providing increased power absorption. 
Aerofoil sections can be employed which feature a rapid drag rise above a 
particular incidence or Mach number, and the rotors could be driven into a 
high drag regime when power absorption is required. A reduction in the 
rotor disc area would increase the power required to produce a given 
thrust and would result in a more compact configuration, and rotor blade 
twist could be selected to promote early tip stall or the progressive 
spread of stall to provide readily controllable power absorption 
characteristics. 
Whilst one embodiment has been described and illustrated, it will be 
understood that various modifications can be made without departing from 
the scope of the invention as defined in the appended claims. Other power 
absorber means could be used to absorb power from the sustaining rotor, 
and in a twin tail rotor configuration, boundary layer control of the tail 
rotor blades could be used to achieve the opposing thrust forces. The 
gearing associated with the yaw demand mechanism may be operated either 
mechanically or electrically.