Note: Descriptions are shown in the official language in which they were submitted.
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This invention relates to aircraft control,
and more particularly, to controlling the powerplant
based on airframe body states to enhance aircraft
maneuvering performance.
In the main hereinafter the control of
helicopters is discussed but the teachings disclosed
herein are relevant to rotorcraft generally.
In modern helicopters, the trend toward main
rotor systems which have lower inertia (angular momen-
tum) reduces the level of stored energy in the rotor
system and causes the rotor to be more susceptible
to large transient speed excursions during some flight
maneuvers. Such main rotor speed excursions, in con-
junction with other flight characteristics of heli-
copters, will change the thrust and control capability
of the rotor and will upset the attitude trim of the
aircraft and cause an undesirable lag in attaining
altitude or speed. An undesirable perturbation of
attitude trim either increases pilot workload (fre-
quently at critical times), saturates the aircraft
stability augmentation system, or both. Therefore,
it is known to provide closed loop fuel control for
controlling rotor speed at a reference speed. Such
a system is disclosed in Canadian Patent 1,202,098,
issued March 18, 1986, entitled FUEL CONTROL FOR
CONTROLLING HELICOPTER ROTOR/TVRBINE ACCELERATION.
However, at times strict control over rotor speed
may be disadvantageous.
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A coordinated turn is equivalent to a pull-up in
terms of loads induced in the helicopter, particularly
main rotor blade loading. This is due to the force
necessarily applied to the helicopter through the
blades in order to effect the required directional
acceleration against the mass of the helicopter and,
in a pull-up, to overcome the acceleration of gravity.
In fact, a 60 bank angle ~which is not uncommon) will
nominally double the loading on the main rotor. Depend-
10 ing on conditions this could cause the rotor to tend to ,
speed up. Since under these conditions torque required
is reducing, it is easily understood that not allowing
the rotor to speed up and demand more torque is counter-
producti~e in such a circumstance. Available rotor
thrust, and hence load factor, could be increased if
rotor speed were allowed to increase.
Consider the following. A helicopter is flying at
cruise speed (e.g., at least sixty knots) and the pilot
initiates a coordinated turn. In one case by virtue of
a combination of control inputs a flight path is chosen
which results in forward speed (and/or altitude) being
allowed to bleed off. Under these conditions, which by
nature of the energy exchange process are transient,
the torque required by the rotor is reduced and a
tendency exists for the rotor to speed up (Kenetic and
or potential energy of the airframe is used up by the
rotor). The existing closed-loop fuel control restrains
this tentency by ba~king down the engine torque to
retain the torque balance between main rotor required
torque and engine supplied torque to preserve the
reference rotor speed, which is undesirable. It would be
desirable in such a circumstance, as taught herein, to
re-reference the rotor speed up, thereby providing the
helicopter with potential for more thrust, from the
` 35- increaset rotor speed and hence the capability to pull
higher levels of load factor. In another case, the pilot
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desires to maintain forward speed (and altitude) in a
steady turn. Under these conditions, in whi~h the
increased thrust (required to maintain the load factor
in the turn) results in a higher level of torque
5 required by the rotor, the engines mNst provide the
energy to maintain closed loop rotor speed control.
Under these circumstances the pilot could, by increasing
- control input, pull increased thrust (and load factor)
up to the power limit of the engines. In a more desirable
10 fashion, as taught herein, installed engine power would
be better utilized by re-referencing the rotor speed to
a higher setting thus preserving a higher stall and
control margin on the rotor. These two specific condi- 3
tions are used for illustration, but there are other
15 levels of maneuvering flight which could benefit from ~,
suitable adjustment of rotor reference speed. Common
to all such maneuvers is the airframe (body) pitch rate
which i8 necessarily generated as part of executing
the maneuver.
20 Disclosure of Invention
Therefore, it is an object of this invention to
overcome the disadvantages of closed-loop rotor speed
control by allowing/causing the rotor to speed up in
a positive load maneuver, thereby increasing the
available thrust, hence allowing ~hi ~ r aircraft load
factors, at cruise speeds. It is a further object to
implement the invention without additional sensors
and with a minimum of additional circuitry where an
AFCS is available.
According to the invention rotor speed, which in
the case of a free turbine engine is directly propor-
tioned to the free turbine speed, is sensed and main-
tained by a closed-loop fuel control at a reference
speed. The reference speed is biased up as a function
of a pitch rate indicative of a positive load maneuver
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to allow/cause the rotor speed to increase in a controlled
-- manner, thereby increasing available thrust and improving
aircraft loading. Z
The invention may be practiced in a variety of
5 analog, digital, or computer controls, in a straight-
forward manner, or with additional features incorporated
therewith to provide a more sophisticated control. The
invention is easily implemented utilizing apparatus
and techniques which are well within the skill of the
10 art, in the light of the specific teachings with
respect thereto which follow hereinafter.
Other objects, features and advantages of the
present invention will become more apparent in the
light of the following detailed description of
15 exemplary embodiments thereof, as illustrated in the
accompanying drawing.
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Brief Description of Drawing
The sole FIGURE herein is a simplified schematic
block diag~am of the fuel control loop of a helicopter
20 incorporating the present invention.
Best Mode for Carrying Out the Invention
In Fig. l is shown a fuel control system for a
helicopter. A main rotor io is connected by a shaft
12 to a gear box 13 which is driven by a shaft 14
25 through an overrunning clutch 16, which engages an
output shaft 18 of an engine 20, but which disengages
during autorotation. The gear box 13 also drives a
tail rotor 22 through a shaft 24 so that the main
rotor 10 and the tail rotor 22 are always turning at
30 speeds having a fixed relationship to each other, such
as the tail rotor rotating about five times faster than
the main rotor.
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The engine 20 as shown comprises a free turbine
gas engine in which the output shaft 18 is driven by
a free turbine 26, which is in turn driven by gases
from a gas generator including a turbocompressor having
a compressor 28 connected by a shaft 30 to a compressor-
driving turbine 32, and a burner section 34 to which
fuel is supplied by fuel lines 36 from a fuel pump 38
through a fuel control metering valve 40.
The fuel control system nominally provides the
correct rate of fuel in the fuel lines 36 so as to
maintain a desired rotor speed. For purposes of this
discussion, autorotation is ignored and free turbine
speed is indicative of rotor speed. Therefore, a
tachometer 42 measures the speed of the free turbine
26 (such as at the output shaft 18) to provide an
actual (rotor) speed signal on a line 44 to a summing
junction 46. Although not referred to herein, the turbine
speed signal on the line 44 may be filtered before appli-
cation to the summing junction 46 in order to eliminate
noise therefrom and to ensure acceptable closed loop
stability margins. A rotor speed reference signal 48,
which typically is set at 100% rated speed, is also
provided to the summing junction 46. The output of the
summing junction 46 is a rotor speed error signal on a
line 52 which is nominally ZER0 or, in other words, the
difference between the actual speed signal and the
reference speed signal. A turbine governor 54 is
responsive to the rotor speed error signal on the line
52 and to the reference signal 48 and, in conjunction
with a gas generator control 58, provides a fuel command
signal to the metering valve 40 so as to cause the
correct amount of fuel from the fuel pump 38 to be
provided in the fuel inlet lines 36 to maintain the
rotor speed at the reference speed. This provides a
servo loop which could be implemented in a number of
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straightforward manners. The rotor speed reference
signal may be biased at the summing junction 46 by
pilot beep commands on a line 50 from a pilot's
engine speed beeper (not shown). The rotor speed
reference signal may also be biased at the summing
junction 46 by a rotor speed reference bias signal
on a line 70. As the rotor speed reference signal
48 is biased (Up), the rotor speed error signal is
driven (biased) from ZERO and the fuel control system
causes the engine (rotor) to be maintained at a higher
reference speed.
With reference to the load factor enhancing
portion of this invention, the induced pitch rate
of the aircraft is sensed by a pitch rate gyro 72
that provides a pitch rate signal which is shaped
by a shaping circuit 74. The shaping circuit 74 may
be embodied in an existing automatic flight control
system (AFCS) 76, and may integrate, amplify, lag,
limit, etc. the pitch rate signal to tailor the rotor
speed increase to the loading needs of a particular
aircraft. (Also, there is typically a rotor speed
above which rotor damage may occur.) The shaping
circuit may be embodied in existing control circuitry,
such as disclosed in U.S. Patent No. 4,127,245 (Tefft,
1978) entitled HELICOPTER PITCH RATE FEEDBACK BIAS
FOR PITCH AXIS MANE~VERING STABILITY AND LOAD FEEL.
(Therein, the signal on the line 32 from the amplifier
34 corresponds to the shaped pitch rate signal des-
cribed herein.) The switch 78 is responsive to an
airspeed signal, provided by an airspeed measuring
means 80, and when closed in response to an airspeed
signal indicative of cruise speed provides the shaped
pitch rate signal to the line 70 as the rotor speed
reference bias signal which references the rotor speed
up, as discussed hereinbefore. The shaping circuit
74 may also be responsive to the airspeed signal,
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for instance to affect the overall sensitivity (gain).
In a like manner, other aircraft parameters could be
sensed to more accurately tailor the response to the
situation.
In a banked turn, even though the pitch attitude
in the inertial axis may remain fixed, a pitch rate is
induced in the body axis (i.e., in a pitch rate gyro
affixed to the helicopter). The induced pitch rate
is proportional to the yaw rate and the sine of the bank
angle. A positive pitch rate ~body axis) maneuver
requires loads in the main rotor in proportion to the
sensed pitch rate to sustain the load factor. The pitch
rate signal is therefore used as an indicator of load
factor to reference the rotor speed up and provide the
potential for increased rotor thrust. For positive
load maneuvers, the rotor speed increases to augment the
level of thrust and subsequent load factor which can be
developed. In one case (i.e., turnin~ with no concern
for forward sDeed/altitude loss). biasin~ the rotor
s~eed reference sivnal com~lements the natural tendencv
for the rotor to s~eed u~. In another case (maintaining
forward speed and altitude while turning), referencing the
rotor speed up provides for potentially higher rotor
thrust while preserving rotor stall and control margins.
Pitch rates indicative of negative load maneuvers are
not used to decrease the rotor reference speed, because
to do so would be undesirable from a control point of
view (among having other complicated side effects).
It should be understood that load factor could be
sensed directly, such as by an accelerometer 73 in the
vertical body axis, to provide a signal that is shaped
to bias the rotor speed reference signal either alone
or in conjunction with the pitch rate signal.
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Although the invention is illustrated in an analog
fashion for clarity, the signal processing functions
involved may preferably be performed in a digital
computer, when one is available. Thus, in a digital
fuel control, the si~nal processing functions of the
invention would be performed by relatively simple
programming steps which are analogous in an obvious
fashion to the signal processing described herein.
Or, a simple hydromechanical gas generator fuel control
capable of receiving a required gas generator speed
signal from the turbine governor 54 could be employed
on a helicopter having a digital automatic flight
control system in which the processing of the engine
speed signal to practice the present invention would
be accomplished by simple programming steps performed
within the automatic flight control computer. But
this is not germane to the inventive concept. It is
sufficient that the invention may be practiced in any
way in which the rotor speed reference signal is biased
as a function of the aircraft pitch rate as sensed by
an on-board pitch rate gyro.
Although the invention has been shown and described
with respect to exemplary embodiments thereof, it
should be understood by those skilled in the art that
the foregoing and various other changes, omissions and
additions may be made therein and thereto, without
departing from the spirit and the scope of the invention.
What is claimed ls:
