GB2144244A

GB2144244A - Rotorcraft load factor enhancer

Info

Publication number: GB2144244A
Application number: GB08417638A
Authority: GB
Inventors: Dean Earl Cooper; James John Howlett
Original assignee: United Technologies Corp
Current assignee: Raytheon Technologies Corp
Priority date: 1983-08-01
Filing date: 1984-07-11
Publication date: 1985-02-27
Also published as: DE3428224A1; DE3428224C2; JPH0733159B2; FR2550161B1; GB8417638D0; GB2144244B; IT8422117A0; JPS6038297A; FR2550161A1; IL72461A; CA1246717A; IT1174616B

Abstract

Rotor speed is maintained at a reference speed 48 by a closed-loop fuel control system 40,44,46,48,54,58. Load factor potential is enhanced by increasing the reference speed as a function of the helicopter pitch rate 72 and optionally airspeed 80 when undergoing positive pitch rate manoeuvres. The reference may be increased only when a certain airspeed (cruise speed) has been reached. Alternatively load factor can be sensed directly or combined with pitch rate by use of an accelerometer 73. <IMAGE>

Description

SPECIFICATION Rotorcraft load factor enhancer Technical field This invention relates to aircraft control, and more particularly, to controlling the powerplant based on airframe body states to enhance aircraft maneuvering performance.

Background art 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 momentum) reduces the level of stored energy in the rotor system and causes the rotor to be more susceptible to large trasient speed excursions during some flight maneuvers. Such main rotor speed excursions, in conjunction with other flight characteristics of helicopters, 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 (frequently 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 copending, commonly-owned U. S. Application No. 369,301, filed on April 16, 1982 entitled FUEL CONTROL FOR CONTROLLING HELICOPTER ROTOR/TURBINE ACCELERATION. However, at times strict control over rotor speed may be disadvantageous.

Acoordinatedturn is equivalentto 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.

Depending 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-productive 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 retrains this tendency by backing 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 increased rotor speed and hence the capability to pull higher levels of load factor. In another case, the pilot desires to maintain forward speed (and altitude) in a steady turn. Under these conditions, in which the increased thrust (required to maintain the load factor in the turn) results in a higher level of torque required by the rotor. the engines must 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 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 conditions are used for illustration, but there are other levels of maneuvering flight which could benefit from suitabie adjustment of rotor reference speed. Common to all such maneuvers is the airframe (body) pitch rate which is necessarily generated as part of executing the maneuver.

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 potentially higher 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 proportioned to the free turbine speed, is sensed and maintained 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 to allow/cause the rotor speed to increase in a controlled manner, thereby increasing available thrust and improving aircraft loading.

The invention may be practiced in a variety of analog, digital, or computer controls, in a straightforward 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 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 exemplary embodiments thereof, as illustrated in the accompanying drawing.

Brief description of drawing The sole FIGURE herein is a simplified schematic block diagram of the fuel control loop of a helicopter incorporating the present invention.

Best mode for carrying out the invention In Figure 1 is shown a fuel control system for a helicopters A main rotor 10 is connected by a shaft 12 to a gear box 13 which is driven by a shaft 14 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 speeds having a fixed relationship to each other, such as the tail rotor rotating about five times faster than the main rotor.

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 suming junction 46. Although not referred to herein, the turbine speed signal on the line 44 may be filtered before application to the summing junction 46 in orderto 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 ZERO or, in otherwords,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 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 show).

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 MANEUVERING STABILITY AND LOAD FEEL, which is incorporated by reference herein. (Therein, the signal on the line 32 from the amplifier 34 corresponds to the shaped pitch rate signal described 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, for instance to affect the overall sensitivity (gain). In a like man i, 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 aug men. the level of thrust and subsequent load factor which can be developed.In one case (i.e., turning with no concern for forward speed/altitude loss) biasing the rotor speed reference signal comple menus the natural tendency for the rotor to speed up.

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.

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 signal 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 is:

Claims

1. A rotorcraft fuel control system, for controlling the rotor speed of the rotorcraft, wherein the rotor is driven by an engine, that comprises: tachometer means for providing an actual speed signal indicative of the actual rotor speed; fuel valve means, responsive to a fuel command signal, for metering fuel to the engine to control the speed of the rotor; and signal processing means, connected for response to the tachometer means and connected to the fuel valve means, for providing a reference speed signal indicative of a desired rotor speed, for providing a rotor speed error signal as the difference of the actual speed signal and the reference speed signal, for providing the fuel command signal to said fuel valve means in response to said rotor speed error signal; characterized by: pitch rate means for providing a pitch rate signal indicative of the pitch rate of the rotorcraft; and said signal processing means connected for response to said pitch rate means and comprising means for providing a rotor reference speed bias signal in response to the pitch rate signal to bias the rotor speed reference signal as a function of said pitch rate signal, thereby biasing the rotor speed error signal and augmenting the fuel command signal in response to pitch rate to increase load factor potential during a positive load maneuver.

2. A rotorcraft fuel control system according to Claim 1 characterized by airspeed measuring means for providing an airspeed signal means indicative of the airspeed of the rotorcraft to said signal processing and wherein said rotor speed reference bias signal is provided as a function of said airspeed signal and said pitch rate signal.

3. A rotorcraft fuel control system according to claim 1 characterized by airspeed measuring means for providing an airspeed signal to said signal processing means indicative of the airspeed of the rotorcraft and wherein said rotor speed reference bias signal is provided only when the airspeed signal is indicative of at least a threshold (cruise) airspeed.

4. A rotorcraft fuel control system, for controlling the rotor speed of the rotorcraft, wherein the rotor is driven by an engine, that comprises: tachometer means for providing an actual speed signal indicative of the actual rotor speed; fuel valve means, responsive to a fuel command signal, for metering fuel to the engine to control the speed of the rotor; and signal processing means, connected for response to the tachometer means and connected to the fuel valve means, for providing a reference speed signal indicative of a desired rotor speed, for providing a rotor speed error signal as the difference of the actual speed signal and the reference speed signal, for providing the fuel command signals to said fuel valve means in response to said rotor speed error signal; characterized by: load factor sensing means for providing a load factor signal indicative of the load factor on the rotorcraft; and said signal processing means connected for response to said load factor sensing means and comprising means for providing a rotor reference speed bias signal in response to the load factor signal to bias the rotor speed reference signal as a function of said load factor signal, thereby biasing the rotor speed error signal and augmenting the fuel command signal in response to load factor to increase load factor potential during a positive load maneuver.

5. A rotorcraft fuel control system according to claim 4 characterized by pitch rate means for providing a pitch rate signal indicative of the pitch rate of the rotorcraftto said signal processing means and wherein said rotor speed reference bias signal is provided as a function of said pitch rate signal and said load factor signal.