CA1246717A - Rotorcraft load factor enhancer - Google Patents
Rotorcraft load factor enhancerInfo
- Publication number
- CA1246717A CA1246717A CA000458431A CA458431A CA1246717A CA 1246717 A CA1246717 A CA 1246717A CA 000458431 A CA000458431 A CA 000458431A CA 458431 A CA458431 A CA 458431A CA 1246717 A CA1246717 A CA 1246717A
- Authority
- CA
- Canada
- Prior art keywords
- rotor speed
- signal
- load factor
- pitch rate
- rotor
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired
Links
- 239000003623 enhancer Substances 0.000 title abstract 2
- 230000001965 increasing effect Effects 0.000 claims abstract description 14
- 230000002708 enhancing effect Effects 0.000 claims description 6
- 230000003750 conditioning effect Effects 0.000 claims 6
- 230000001143 conditioned effect Effects 0.000 claims 3
- 239000000446 fuel Substances 0.000 abstract description 21
- 239000007789 gas Substances 0.000 description 6
- 238000007493 shaping process Methods 0.000 description 4
- RZVHIXYEVGDQDX-UHFFFAOYSA-N 9,10-anthraquinone Chemical compound C1=CC=C2C(=O)C3=CC=CC=C3C(=O)C2=C1 RZVHIXYEVGDQDX-UHFFFAOYSA-N 0.000 description 3
- 230000001133 acceleration Effects 0.000 description 3
- 230000008901 benefit Effects 0.000 description 2
- 230000000694 effects Effects 0.000 description 2
- 238000000034 method Methods 0.000 description 2
- 230000004044 response Effects 0.000 description 2
- 230000001052 transient effect Effects 0.000 description 2
- RZVAJINKPMORJF-UHFFFAOYSA-N Acetaminophen Chemical compound CC(=O)NC1=CC=C(O)C=C1 RZVAJINKPMORJF-UHFFFAOYSA-N 0.000 description 1
- 238000007792 addition Methods 0.000 description 1
- 230000003416 augmentation Effects 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- 229910000078 germane Inorganic materials 0.000 description 1
- 230000005484 gravity Effects 0.000 description 1
- 238000005381 potential energy Methods 0.000 description 1
- 230000008569 process Effects 0.000 description 1
- 230000035945 sensitivity Effects 0.000 description 1
Classifications
-
- G—PHYSICS
- G05—CONTROLLING; REGULATING
- G05D—SYSTEMS FOR CONTROLLING OR REGULATING NON-ELECTRIC VARIABLES
- G05D1/00—Control of position, course or altitude of land, water, air, or space vehicles, e.g. automatic pilot
- G05D1/08—Control of attitude, i.e. control of roll, pitch, or yaw
- G05D1/0808—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft
- G05D1/0858—Control of attitude, i.e. control of roll, pitch, or yaw specially adapted for aircraft specially adapted for vertical take-off of aircraft
-
- H—ELECTRICITY
- H02—GENERATION; CONVERSION OR DISTRIBUTION OF ELECTRIC POWER
- H02P—CONTROL OR REGULATION OF ELECTRIC MOTORS, ELECTRIC GENERATORS OR DYNAMO-ELECTRIC CONVERTERS; CONTROLLING TRANSFORMERS, REACTORS OR CHOKE COILS
- H02P23/00—Arrangements or methods for the control of AC motors characterised by a control method other than vector control
- H02P23/16—Controlling the angular speed of one shaft
Abstract
Rotorcraft Load Factor Enhancer Abstract Rotor speed is maintained at a reference speed by a closed-loop fuel control system. Load factor potential is enhanced by increasing the reference speed as a function of the helicopter pitch rate when undergoing positive pitch rate maneuvers.
Description
-" ~2467i7 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.
~ . _ '.~ ,~.
12g~717 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 ^
~246717 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 --` . 1246717 , 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.
~: ;
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.
, , .. .
.
~.
- ` ~246717 _ 5 " ~
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 ~ ~246717 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, '--` , 12g6~17 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.
~ 12467~7 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:
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.
~ . _ '.~ ,~.
12g~717 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 ^
~246717 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 --` . 1246717 , 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.
~: ;
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.
, , .. .
.
~.
- ` ~246717 _ 5 " ~
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 ~ ~246717 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, '--` , 12g6~17 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.
~ 12467~7 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:
Claims (5)
1. Load factor enhancing apparatus for a heli-copter including a rotor driven by an engine, comprising:
means for providing a rotor speed reference signal indicative of a desired rotor speed;
means responsive to the rotor speed reference signal for controlling the rotor speed at the desired rotor speed;
characterized by:
pitch rate means for providing a pitch rate signal indicative of the pitch rate of the helicopter in a positive load maneuver;
means for conditioning the pitch rate signal according to an increased rotor speed that meets the loading requirements of the helicopter;
and means for adding the conditioned pitch rate signal to the rotor speed reference signal so as to increase rotor speed above the desired rotor speed to the increased rotor speed during the positive load maneuver, thereby increasing load factor potential in the maneuver.
means for providing a rotor speed reference signal indicative of a desired rotor speed;
means responsive to the rotor speed reference signal for controlling the rotor speed at the desired rotor speed;
characterized by:
pitch rate means for providing a pitch rate signal indicative of the pitch rate of the helicopter in a positive load maneuver;
means for conditioning the pitch rate signal according to an increased rotor speed that meets the loading requirements of the helicopter;
and means for adding the conditioned pitch rate signal to the rotor speed reference signal so as to increase rotor speed above the desired rotor speed to the increased rotor speed during the positive load maneuver, thereby increasing load factor potential in the maneuver.
2. Load factor enhancing apparatus according to claim 1 characterized by:
airspeed means for providing an airspeed signal indicative of the airspeed of the helicopter;
wherein the means for conditioning the pitch rate signal is responsive to the airspeed signal and the conditioning is a function thereof.
airspeed means for providing an airspeed signal indicative of the airspeed of the helicopter;
wherein the means for conditioning the pitch rate signal is responsive to the airspeed signal and the conditioning is a function thereof.
3. Load factor enhancing apparatus according to claim 1 characterized by:
airspeed means for providing an airspeed signal indicative of the airspeed of the helicopter;
means for permitting the addition of the conditioned pitch rate signal to the rotor speed reference signal only when the airspeed signal is indicative of at least a threshold airspeed.
airspeed means for providing an airspeed signal indicative of the airspeed of the helicopter;
means for permitting the addition of the conditioned pitch rate signal to the rotor speed reference signal only when the airspeed signal is indicative of at least a threshold airspeed.
4. Load factor enhancing apparatus for a helicopter including a rotor driven by an engine, comprising:
means for providing a rotor speed reference signal indicative of a desired rotor speed;
means responsive to the rotor speed refer-ence signal for controlling the rotor speed at the desired rotor speed;
characterized by:
load factor means for providing a load factor signal indicative of the load factor of the helicopter in a positive load maneuver;
means for conditioning the load factor signal according to an increased rotor speed that meets the loading requirements of the helicopter; and means for adding the conditioned load factor signal to the rotor speed reference signal so as to increase rotor speed above the desired rotor speed to the increased rotor speed during the positive load maneuver, thereby increasing load factor potential in the maneuver.
means for providing a rotor speed reference signal indicative of a desired rotor speed;
means responsive to the rotor speed refer-ence signal for controlling the rotor speed at the desired rotor speed;
characterized by:
load factor means for providing a load factor signal indicative of the load factor of the helicopter in a positive load maneuver;
means for conditioning the load factor signal according to an increased rotor speed that meets the loading requirements of the helicopter; and means for adding the conditioned load factor signal to the rotor speed reference signal so as to increase rotor speed above the desired rotor speed to the increased rotor speed during the positive load maneuver, thereby increasing load factor potential in the maneuver.
5. Load factor enhancing apparatus according to claim 1 characterized by:
pitch rate means for providing a pitch rate signal indicative of the pitch rate of the helicopter;
wherein the means for conditioning the load factor signal is responsive to the pitch rate signal and the conditioning is a function thereof.
pitch rate means for providing a pitch rate signal indicative of the pitch rate of the helicopter;
wherein the means for conditioning the load factor signal is responsive to the pitch rate signal and the conditioning is a function thereof.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US51933283A | 1983-08-01 | 1983-08-01 | |
| US519,332 | 1983-08-01 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1246717A true CA1246717A (en) | 1988-12-13 |
Family
ID=24067852
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000458431A Expired CA1246717A (en) | 1983-08-01 | 1984-07-09 | Rotorcraft load factor enhancer |
Country Status (7)
| Country | Link |
|---|---|
| JP (1) | JPH0733159B2 (en) |
| CA (1) | CA1246717A (en) |
| DE (1) | DE3428224C2 (en) |
| FR (1) | FR2550161B1 (en) |
| GB (1) | GB2144244B (en) |
| IL (1) | IL72461A (en) |
| IT (1) | IT1174616B (en) |
Families Citing this family (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB8807676D0 (en) * | 1988-03-31 | 1988-05-05 | Westland Helicopters | Helicopter control systems |
| US4998202A (en) * | 1989-05-19 | 1991-03-05 | United Technologies Corporation | Helicopter, high rotor load speed enhancement |
| US5314147A (en) * | 1991-08-27 | 1994-05-24 | United Technologies Corporation | Helicopter engine speed enhancement during heavy rotor load and rapid descent rate maneuvering |
| US5265826A (en) * | 1991-08-27 | 1993-11-30 | United Technologies Corporation | Helicopter engine control having lateral cyclic pitch anticipation |
| US5265825A (en) * | 1991-08-27 | 1993-11-30 | United Technologies Corporation | Helicopter engine control having yaw input anticipation |
| FR3000465B1 (en) | 2012-12-27 | 2015-02-13 | Eurocopter France | METHOD FOR ROTATING A MAIN ROTOR OF ROTOR OF ROTOR, ACCORDING TO A VARIABLE VALUE ROTATION SPEED SET |
| FR3000466B1 (en) | 2012-12-27 | 2015-02-13 | Eurocopter France | METHOD FOR ROTATING A ROTOR OF A ROTOR BY FORECKING ANTICIPATION OF TORQUE REQUIREMENTS BETWEEN TWO ROTATOR ROTATION SPEED INSTRUCTIONS |
Family Cites Families (14)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2645293A (en) * | 1946-09-14 | 1953-07-14 | Gen Electric | Apparatus for regulating propulsion power plants |
| DE952404C (en) * | 1953-05-18 | 1956-11-15 | Bendix Aviat Corp | Automatic control equipment for aircraft |
| US3174551A (en) * | 1963-02-19 | 1965-03-23 | United Aircraft Corp | Power management control for helicopters |
| US3200886A (en) * | 1964-05-28 | 1965-08-17 | Joseph L Magri | Droop compensated fuel control system |
| GB1120327A (en) * | 1967-02-24 | 1968-07-17 | Ltv Aerospace Corp | Autothrottle |
| GB1244160A (en) * | 1967-11-17 | 1971-08-25 | Dowty Rotol Ltd | Engine, propeller and rotor installations |
| US3930366A (en) * | 1974-07-17 | 1976-01-06 | General Motors Corporation | Helicopter power plant control |
| JPS5112012A (en) * | 1974-07-18 | 1976-01-30 | Mitsubishi Heavy Ind Ltd | GASUTAABINNENSHOKINO KAENKANSHISOCHI |
| US4217754A (en) * | 1977-01-22 | 1980-08-19 | Bodenseewerk Geratetechnik Gmbh | Apparatus for controlling the rotary speed in turbo-jet engines for aircraft |
| US4127245A (en) * | 1977-04-27 | 1978-11-28 | United Technologies Corporation | Helicopter pitch rate feedback bias for pitch axis maneuvering stability and load feel |
| GB2052805B (en) * | 1979-06-29 | 1983-03-09 | Smiths Industries Ltd | Gas-turbine engine control |
| ZA814691B (en) * | 1980-08-08 | 1983-02-23 | Ass Eng Ltd | Automatic speed control systems |
| US4442667A (en) * | 1981-01-14 | 1984-04-17 | Aviation Electric Ltd. | Acceleration limit reset |
| US4423593A (en) * | 1982-04-16 | 1984-01-03 | Chandler Evans Inc. | Fuel control for controlling helicopter rotor/turbine acceleration |
-
1984
- 1984-07-09 CA CA000458431A patent/CA1246717A/en not_active Expired
- 1984-07-11 GB GB08417638A patent/GB2144244B/en not_active Expired
- 1984-07-19 JP JP59150471A patent/JPH0733159B2/en not_active Expired - Fee Related
- 1984-07-20 IL IL7246184A patent/IL72461A/en not_active IP Right Cessation
- 1984-07-30 IT IT22117/84A patent/IT1174616B/en active
- 1984-07-30 FR FR8412065A patent/FR2550161B1/en not_active Expired
- 1984-07-31 DE DE3428224A patent/DE3428224C2/en not_active Expired - Lifetime
Also Published As
| Publication number | Publication date |
|---|---|
| GB2144244B (en) | 1986-11-12 |
| DE3428224A1 (en) | 1985-02-14 |
| FR2550161A1 (en) | 1985-02-08 |
| IT1174616B (en) | 1987-07-01 |
| IL72461A (en) | 1996-10-16 |
| GB8417638D0 (en) | 1984-08-15 |
| FR2550161B1 (en) | 1988-06-10 |
| GB2144244A (en) | 1985-02-27 |
| JPH0733159B2 (en) | 1995-04-12 |
| DE3428224C2 (en) | 2000-02-10 |
| JPS6038297A (en) | 1985-02-27 |
| IT8422117A0 (en) | 1984-07-30 |
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| Date | Code | Title | Description |
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| MKEX | Expiry |