JPH0733159B2 - Fuel controller for rotorcraft - Google Patents

Description

Description: BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to an aircraft control system, and more particularly, to a fuel control device that controls a power unit based on a state of an airframe to enhance the dynamic performance of a rotary wing aircraft. Pertain to.

2. Description of the Related Art The following mainly describes control of a helicopter, but the content disclosed herein is generally applied to a rotary-wing aircraft.

In recent helicopters, the inertia of the main rotor and its drive system is relatively small due to the weight reduction of the main rotor and its drive system, and the energy stored in the main rotor and its drive system is correspondingly small. There is a tendency for large speed fluctuations to occur easily. Such speed fluctuations of the main rotor, coupled with other flight characteristics of the helicopter,
Changing the thrust of the rotor and its adaptability to control,
It is undesirable in that it disturbs the attitude control of the aircraft and causes a time delay in obtaining a predetermined altitude or speed. Disturbance in attitude control increases the work load on the pilot, especially in an emergency, and may cause an insufficient operation even if the stability increasing device of the aircraft operates to its full operating capacity, which is not preferable. Therefore, it is known to perform closed loop fuel control so that the rotor speed is controlled based on a reference value. Such a system is disclosed in Japanese Patent Application No. 58-066825 (Japanese Patent Publication No. 4-16618), which is filed by the same applicant as the present applicant.
No.).

When attempting to raise the aircraft, of course, the output of the main rotor must be increased, and a control command for increasing the output for that purpose is given by the pilot's maneuvering motion, and a response to the control command is given. The result is clearly fed back to the control commander, that is, the pilot as a change in flight altitude. On the other hand, even if the altitude of the machine body is constant, the tilted rotation of the machine body corresponds to a sudden rise of the machine body from the viewpoint of the load applied to the main rotor. This is because a tilting turn requires a force that produces an acceleration to orient the aircraft in the required direction against the centrifugal force that acts on the mass of the helicopter. Similar to requiring power to overcome acceleration. In fact, a 60 ° tilted turn, which is by no means anomalous, usually doubles the load on the main rotor. During such a tilted turn, the kinetic energy of the airframe tends to drive the rotor, conversely, by air, reducing the torque required by the rotor and increasing the speed of the rotor. In this case, according to the existing closed loop fuel control system, the engine torque is reduced to keep the rotor speed at the reference value so that the torque balance between the torque required by the main rotor and the torque supplied by the engine is maintained. It tries to suppress the tendency of the rotor speed to increase, but such control is undesirable and
Further, the change in the flight path caused by the lack of the input torque to the main rotor appears as a combination of the change in the altitude of the body and the change in the horizontal position, so that the feedback to the pilot is not accurately performed. In such a state, it is desirable to set the reference value of the rotor speed to a higher value, and based on that, allow a larger rotor speed to the helicopter, make its thrust generation capacity higher, and set the load multiple to a higher level. Is to give the ability to pull up.

Accordingly, it is an object of the present invention to automatically increase the rotor speed during a load-increasing flight such that an increase in load on the rotor cannot be accurately determined by the pilot, such as in tilting turns during cruise, thereby increasing rotor speed. In a way that increases thrust, and thus enhances the load multiple capacity of the aircraft,
It is an object of the present invention to provide a fuel controller for a rotary wing aircraft that makes up for the drawbacks of closed loop rotor speed control.

According to the present invention, there is provided a fuel control device for a rotary wing aircraft including a rotor and an engine for driving the rotor, wherein the engine control unit responds to a fuel command signal. A fuel control valve for metering the fuel supplied to the engine to control the output; means for producing a first signal for controlling the output of the engine in response to pilot maneuvering movement; and at least rotation of the rotor. Means for producing a flight parameter signal including a second signal representative of velocity and a third signal representative of a pitch rate of the aircraft based on an inertial coordinate system; and responsive to at least the first, second and third signals. Controlling the fuel supply to the engine according to the difference between the first signal and the second signal and increasing the output of the engine as the value of the third signal increases. Is achieved by a fuel control apparatus for a rotary wing aircraft comprising a signal processing means for biasing said fuel command signal in response to the magnitude of the third signal.

According to the present invention, the rotor speed (in a free turbine engine, which is directly proportional to the free turbine speed) is detected, and the reference value of the rotor speed is maintained by the closed loop fuel control. In a fuel control device of a type aircraft, such a reference value is increased as a function of a pitch rate (a change speed of a body pitch angle) which is one of parameters related to a load multiple, and an actual value is increased according to the increase of the pitch rate. By controlling the rotor speed on the basis of a larger reference value, the load characteristics during an inclined turn of the aircraft are improved.

Parameters relating to the load multiple include, in addition to the pitch rate signal, an acceleration signal by an accelerometer along the vertical axis of the machine body, etc. Of course, these other parameter signals may be used in addition to the control based on the pitch rate. . For even finer control, airspeed may also be used with the pitch rate as a parameter to increase the rotor speed reference.

The invention may be implemented with various analog, digital or computer controls, in a simple form, or incorporating additional features to provide a higher degree of control. The present invention is readily implemented using equipment and techniques familiar to one of ordinary skill in the art based on the disclosure below. The present invention can do the above with minimal additional circuitry without additional sensors when utilizing AFCS.

The objects, features and advantages of the present invention will become more apparent in the following detailed description of the exemplary embodiments shown in the drawings.

Examples The accompanying figures show a fuel control system for a helicopter. The main rotor 10 is connected by a shaft 12 to a gearbox 13, which is driven by a shaft 14 through an overrunning clutch 16. clutch
16 is engaged with the output shaft 18 of the engine 20, but is disengaged during autorotation in which the main rotor is driven by the airflow and rotates faster than the speed driven by the engine. . The gearbox 13 also has a shaft 24 so that the main rotor and the tail rotor 22 always rotate at a constant speed, for example, the tail rotor rotates 5 times faster than the main rotor.
The tail rotor 22 is driven through.

As shown, the engine 20 is a free turbine
It includes a gas engine in which the output shaft 18 is driven by a free turbine 26, which is driven by gas from a gas generator. Gas generator shaft 30
Compressor coupled to compressor drive turbine 32 by
It includes a turbocompressor having 28 and a burner section 34 which is fueled by a fuel line 36 from a fuel pump 38 through a fuel control metering valve.

The fuel control system will normally provide the correct amount of fuel in the fuel tube 36 to maintain the desired rotor speed. now,
Ignoring autorotation, the free turbine speed shall indicate rotor speed. Therefore, the tachometer 42
Provides a speed signal on conductor 44 to summing point 46 by measuring the speed of free turbine 26 (eg, the speed at output shaft 18) indicating the actual rotor speed. Although not shown here, the turbine speed signal on conductor 44 is filtered through a suitable filter before it is applied to summing point 46 to remove noise and to ensure an acceptable closed loop stability margin. May be.

A reference signal 48, which is typically set at 100% rated speed and provides a target value for rotor speed, is also provided at summing point 46. Score 46
Is the error signal provided on conductor 52 which gives the difference between the speed signal representing the actual rotor speed and the reference signal representing the target speed, which is zero in the scale.

The turbine governor 54 is responsive to the error signal on line 52 and the reference signal 48 and, in combination with the gas generator control portion 58, delivers the correct amount of fuel from the fuel pump 38 to maintain the rotor speed at the reference value. A fuel command signal is provided to metering valve 40 to supply fuel inlet tube 36. This forms a servo loop that can be implemented in many simple forms.

Reference signal 48 is biased at summing point 46 by a pilot beep command on conductor 50 from a pilot operated engine speed change beeper (not shown). The reference signal is also biased at summing point 46 by the bias signal on conductor 70. At the summing point, when the reference signal 48 is biased upwards, the error signal is biased from 0 to the positive side and the fuel control system controls the engine (and rotor) relative to the higher reference value.

Next, the part relating to the increase in load multiple of the present invention will be described. The pitch rate gyro 72 detects the pitch rate induced in the aircraft. As is well known, a gyro device is a device that can maintain a constant inertial coordinate system without being affected by the attitude of an aircraft body, and the pitch rate detected by such a device is naturally a pitch rate with respect to the inertial coordinate system. Therefore, when the aircraft makes an inclined turn, the increase rate of the pitch of the airframe caused by the inclined turn is detected as compared with the straight flight. The detected pitch rate signal is shaped by the shaping circuit 74. The shaping circuit 74 may be included in an existing automated flight control system (AFCS) and integrates the pitch rate signal to adjust the rotor speed increment appropriately to meet the demands of a particular aircraft load.
Amplification, delay, limiting, etc. may be performed. (Also typically, rotor speed limits are set above which rotor damage can occur.) Shaping circuits are described in US Pat.
It may be implemented in an existing control circuit as disclosed in 127,245 (1978). (Here, the signal on conductor 32 from amplifier 34 corresponds to the shaped pitch rate signal described herein.) Switch 78 is responsive to the airspeed signal provided by airspeed measuring means 80. , When the switch 78 is closed in response to the airspeed signal indicating a certain threshold value, for example, cruise speed, a pitch rate signal shaped as a bias signal which increases the reference value of the rotor speed as described above is provided on the conductor. Given to 70.

The shaping circuit 74 may be adapted to adjust the overall sensitivity (gain) in response to an airspeed signal, for example. Similarly, other aircraft parameters may be detected and utilized to provide a response that more accurately matches the situation.

In the lateral tilt turn, the pitch attitude can be constant with respect to the earth, but the pitch rate is induced in the pitch rate gyro fixed inside the body. The induced pitch rate is proportional to the yaw rate (pivoting angular velocity) and the sine of the lateral tilt angle. A flight with a positive pitch rate
To keep the load multiple constant, a load proportional to the pitch rate is required on the main rotor. Therefore, the pitch rate signal can be used as an indicator of the load multiple, and the control that increases the reference value of the rotor speed according to its positive value and gives a larger rotor thrust is obtained. In load-bearing flights, rotor speed increases, increasing thrust levels and therefore load multiples.

In one case, for example, for turns that are free of forward speed and altitude loss, the reference signal may be slightly increased to provide control that matches the natural tendency of the rotor speed to increase. For turns that maintain forward speed and altitude, the reference signal is increased correspondingly to obtain a higher rotor thrust generation capability while preserving the rotor stall margin and control margin.

Even if a pitch rate occurs which indicates a flight in which the load on the rotor is reduced, this is not used to reduce the reference speed of the rotor. This is because doing so is undesired from a control point of view, especially in that other complicated side effects occur. The load multiple may be detected and shaped directly by an accelerometer 73 or the like along the vertical axis of the airframe, alone or in combination with the pitch rate signal to bias the rotor speed reference signal.

In the above, the invention has been described in analog form for ease of understanding, but if a digital computer is available, the signal processing functions included in the invention are implemented within the digital computer. Is advantageous. Within the digital fuel control system, the signal processing functions of the present invention are implemented by relatively simple program steps similar to the signal processing described herein.

Alternatively, a simple hydromechanical gas generator fuel controller capable of receiving the required gas generator velocity signal from the turbine governor 54 is used in a helicopter with a digital autopilot (AFCS). May be. Within AFCS, processing of engine speed signals to implement the present invention may be accomplished by simple program steps. However, this is not directly related to the technical idea of the present invention.

The essence of the invention is that the rotor speed reference signal is biased as a function of a parameter indicative of the load multiple, for example the aircraft pitch rate detected by an onboard pitch rate gyro, which can be any method. Can be implemented in.

While the present invention has been shown and described with respect to exemplary embodiments thereof, it will be understood by those skilled in the art that various changes, omissions and additions may be made within the scope of the present invention. Will be done.

[Brief description of drawings]

The drawing is a simplified block diagram of a helicopter fuel control loop incorporating the present invention. 10 …… Main rotor, 13 …… Gearbox, 14 …… Axis, 16 ……
Overrunning Kratte, 18 …… shaft, 20 …… engine, 22 …… tail rotor, 24 …… shaft, 26 …… free turbine, 28 …… compressor, 30 …… shaft, 32 …… compressor drive Turbine, 34 …… Burner part, 36 …… Fuel pipe, 38 …… Fuel valve, 4
0 …… Fuel control metering valve, 42 …… Tachometer, 46 …… Addition point,
48 …… 100% reference value, 54 …… Turbine governor, 58 …… Gas generator controller, 72 …… Pitch rate gyro, 74 …… Shaping circuit, 76 …… Automatic flight control system, 78 …… Switch, 8
0 …… Airspeed detection means

Claims (1)

[Claims] 1. A fuel control device for a rotary wing aircraft comprising a rotor and an engine for driving the rotor, wherein fuel supplied to the engine is controlled to control an output of the engine in response to a fuel command signal. A fuel control valve for metering, means for producing a first signal for controlling the output of the engine in response to pilot piloting operation, at least a second signal representing the rotational speed of the rotor and an inertial coordinate system Means for producing a flight parameter signal including a third signal representative of a pitch rate of the aircraft based on the first signal, the second signal and the at least first, second and third signals. The fuel command signal is controlled in accordance with the magnitude of the third signal so that the fuel supply to the engine is controlled according to the difference between the third signal and the higher the value of the third signal, the higher the output of the engine. A fuel control device for a rotary wing aircraft, comprising: