CA1172730A - Fuel control system for gas turbine engine - Google Patents
Fuel control system for gas turbine engineInfo
- Publication number
- CA1172730A CA1172730A CA000391111A CA391111A CA1172730A CA 1172730 A CA1172730 A CA 1172730A CA 000391111 A CA000391111 A CA 000391111A CA 391111 A CA391111 A CA 391111A CA 1172730 A CA1172730 A CA 1172730A
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- Canada
- Prior art keywords
- signal
- engine
- control system
- speed
- fuel control
- 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
- 239000000446 fuel Substances 0.000 title claims abstract description 78
- 230000010354 integration Effects 0.000 claims abstract description 4
- 230000001133 acceleration Effects 0.000 claims description 47
- 230000008859 change Effects 0.000 description 13
- 230000004044 response Effects 0.000 description 12
- 230000001052 transient effect Effects 0.000 description 8
- 230000000454 anti-cipatory effect Effects 0.000 description 4
- 230000007423 decrease Effects 0.000 description 4
- 230000001105 regulatory effect Effects 0.000 description 4
- 230000001276 controlling effect Effects 0.000 description 3
- 208000036366 Sensation of pressure Diseases 0.000 description 2
- 230000003247 decreasing effect Effects 0.000 description 2
- 239000000203 mixture Substances 0.000 description 2
- 230000003534 oscillatory effect Effects 0.000 description 2
- 230000003466 anti-cipated effect Effects 0.000 description 1
- 230000008901 benefit Effects 0.000 description 1
- 230000005540 biological transmission Effects 0.000 description 1
- 230000001419 dependent effect Effects 0.000 description 1
- RLQJEEJISHYWON-UHFFFAOYSA-N flonicamid Chemical compound FC(F)(F)C1=CC=NC=C1C(=O)NCC#N RLQJEEJISHYWON-UHFFFAOYSA-N 0.000 description 1
- 230000006872 improvement Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- 238000005507 spraying Methods 0.000 description 1
- 230000007704 transition Effects 0.000 description 1
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C9/00—Controlling gas-turbine plants; Controlling fuel supply in air- breathing jet-propulsion plants
- F02C9/26—Control of fuel supply
- F02C9/28—Regulating systems responsive to plant or ambient parameters, e.g. temperature, pressure, rotor speed
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2270/00—Control
- F05D2270/01—Purpose of the control system
- F05D2270/02—Purpose of the control system to control rotational speed (n)
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Control Of Turbines (AREA)
Abstract
ABSTRACT
A fuel control system (11) for a gas turbine engine (10) is disclosed. The proportional control regulates the fuel flow Wf from a hydromechanical unit (36) based upon an error signal E representing the difference between a demanded speed signal Nd and an actual speed signal Ng. The demarded speed signal Nd is generated by an integrator (32) as the time integral of the difference between the demanded signal Nd and a scheduled speed signal Ngs. Additionally, the difference signal is limited between an upper limit Al and a lower limit Dl prior to the integration by a comparator (28) and limiter circuit (30).
(Figure 1)
A fuel control system (11) for a gas turbine engine (10) is disclosed. The proportional control regulates the fuel flow Wf from a hydromechanical unit (36) based upon an error signal E representing the difference between a demanded speed signal Nd and an actual speed signal Ng. The demarded speed signal Nd is generated by an integrator (32) as the time integral of the difference between the demanded signal Nd and a scheduled speed signal Ngs. Additionally, the difference signal is limited between an upper limit Al and a lower limit Dl prior to the integration by a comparator (28) and limiter circuit (30).
(Figure 1)
Description
7 ~ ~ ~/a .~
~273~
FUEL CONTROL SYSTEM FOR GAS TURBINE ENGINE
BACKGROUND OF THE INVENTION
-- . . ;
The invention relates to a fuel control system for a gas turbine engine.
For the improved fuel control of gas turbine engines it is known to use a closed loop govenor based on the speed of the turbine rotor. Proportionai governing is introduced in these controls by differencing the actual speed of the turbine and a demandec' or scheduled speed. The error signal developed by differencing these parameters is multiplied b~/ a gain and thereafter used to control fuel flow to change turbine speed in a direction tending to null the error. A pure pro-portional type of cortrol, however~ produces a steady state or perma-nent error between the actual and desTred speeds which increases with turbine speedO This permanent error or "droop" is indicative of the steady state fuel flow at the quiescent operating points of the control-ler. Although the error is inversely proportional to the gain of the proportional loop, it can only be truly nulled with a loop of infinite gain. However, Increasing the gain of a proportional loop to where the "droop" error becomes acceptable can result in instability and loss of control of speed for the gas turbine engine.
To overcome this~ problem, isochronous governors for gas turbine engines have been proposed where the permanent error signal is integrated to zero without producing instability in the proportional loop. The deviation or "droop" error between the scheduled and actual speed signals can then be effectively nulled in this manner. Implemen-tation of tha techni~ue comprises an integral loop in parallel with he proportional loop. The integral loop includes an integrator for ~enerating the time integral of the difference between the scheduled speed and the actual speed. In this manner the integrator output yields a datum to which the proportional loop can be slaved by differencing ;t with the actual speed. Thus, it is the difference between the datum and the actual speed which generates the permanent error signal and steady state fuel control value. In this type o~ system the actual speed becomes slaved to the datum and changes proportionally according to the loop gain with changes in the datum.
~ ~7273~
When the controlling parameter of a slave-datum system is a scheduled speed, the deviation input to the integral loop is an acceleration term since the time integral of acceleration is velocity.
The system is closed loop on acceleration where the magnitude of the difference between the scheduled speed and the actual speed is indica-tive of a scheduled acceleration. The actual acceleration of the engine is fed bac~ through the changes in the actual speed signal to close the loop. It is evident that acceleration limiting of the gas turbine engine can be accomplished in a facile manner by regulating the accelera-tion term before it is input to the integrator~
Although this type of isochronous system provides the advan-tages of being able to null the droop of the governor and limit accelera-tion it does have one disadvantage. The system response to changing transient conditions is regulated by the response of the gas turbine engine to thcse conditions. Before the error caused by a change in operating condttions or a scheduled acceleration can be integrated to a nullity9 the engine must actually change speed an amount equal to the error. Therefore, changes in the datum to which the proportional loop is slaved are dependent upon the inertial lag of the engine. It would be highly desirable to remove this inertial time delay from the integral loop.
It has now been determined that the actual turbine speed will proportionally follow the demanded speecl sign31 or datum by the inertial lag of the engine and thus it can be assumed that the demanded speed signal contains useful information about the anticipated rate of change of actual turbine speed. This assumption is reliably based on the premise that the engine response characteristic is known and it will change speed as commanded by the demanded speed signal. An anticipatory estimate of that rate of change of actual speed can be obtained by comparing the scheduled speed signal to the demanded speed signal ~o generate the acceleration term, thus removing the engine lag from the integral loop calculation.
SUMMARY OF THE INVENTION
The invention is an improved closed loop fuel control system for gas turbine engines. The invention comprises a proportional loop ~7~7~
which differences a demanded speed signal (datum) and an actual speed signa1 and an integra1 loop which integrates an acceleration term to yield the demanded speed signal. The proportional loop is slaved to the demanded speed signal to generate an error signal which controls fuel flow of the engine in a direction tending to null the acceleration term. The acceleration term is integrated to zero for steady state con-ditions by the integral loop. The acceleration term instead of being formed in the normal manner is formed by differencing a scheduled speed signal and the demanded speed signal. This system is closed loop on acceleration because the rate of change of the demar,ded speed signal is utilized as an anticipatory estimate of actual acceleration. Therefore, the acceleration term will lead the actual acce1eration as the demanded speed signa1 1eads the actua1 speed signa1.
By comparing the scheduled speed signa1 to the demanded speed signa1 a higher proportiona1 loop gain may be used in the gas turbine engine fue1 contro1 without producing instabi1ity. This is because the demanded signa1 anticipates the steady state va1ue to which the engine is acce1erating and does not overshoot this va1ue.
Therefore, the actua1 engine speed which proportionally fol10ws the demanded speed signa1 by the inertia1 la~ of the engine will also begin dece1erating before it reaches Its steady state va1ue. The gain of the proportiona1 100p can be increased because the datum is reduced automatica11y upon nearing the steady state value and will not cause the slaved actual speed to overshoot the value.
In a preferred embodiment the improved fuel control is used in conjunction with a hydromechanical unit that regulates fuel flow proportiona11y in dependence on the error signa1. The error signa1 is generated by the contro1 as a function of Wf~Pc where Wf is fuel f10w and Pc is compressor pressure. The hydromechanical unit receives `the error signal in term of the fuel~air ratio and mechanical1y multi-p1ies the signal by a parameter based on compressor pressure to modu-1ate fue1 flow to the engine.
In the preferred embodiment the acce1eration term is also 1imited prior to integration by an acce1eration and dece1eration 1imit based upon the surge 1ine of the particu1ar engineO This configuration ~L72~30 is particularly advantageous in that for transient response the integrator can be set on limit until the acceleration term is reduced to below the limit, whereafter the demanded speed signal can smoothly approach the desired steady state value in a controlled manner. In spite of wherever the original transient begins or its size, this fea~ure produces the same final behavior from the engine. Thus, the fina1 transition from the acceleration schedule to a steady state value can be tailored to bring the system into equilibrium in an optimum amount of time without overshoot.
Another feature of the invention provides that the scheduled speed term is generated by a speed scheduling circuit as a function of the demanded output power of the gas turbine engine. In one pre-ferred implementation the speed scheduling circuit is a frea turbine speed governor for a shaft turbine engine. In a second preferred implementation the speed scheduling circuit governs the substantially constant rotor speed of a helicopter rotor with respect to a reference speed. The output power of the shaft turbine engine in the second implementation is controlled by the operator as a function of the collective pitch angle of the rotor.
These and other objects, features, and aspects of the inven-tion will be better understood and more clearly explained if a reading of the detailed disclosure is undertaken in conjunction with the appended drawings, wherein:
BRIEF DESCRIPTION OF THE DRAWINGS
Figure 1 is a system block diagrammatic vtew of a fuel control system for a turhojet engine which is constructed in accordance with the teachings of the invention; and Figure 2 is a block diagrammatic view of the fuel control system illustrated in Figure 1 in an embodiment suitable for a shaft 3 turbine engine;
Figure 3 is a graphical representation of the steady state values of the demanded speed signal and the actual speed signal as a function of quiescent turbine speed for the fuel control system illus-trated in Figure l; and ~7~730 Figure 4 is a time trace comparison of the demaned speed signal and actual speed signal for the system disclosed in Figure 1 and a prior art system.
DETAILED DESCRIPTION OF THE PREFER~ED EMBODIMENTS
.. . ..
With reference to Figure 1 there is shown a fuel control system 11 for a turbojet engine which is generally designated 10. The engine 10 includes an intake duct 2~ through which air enters. After entering the engine7 the air is compressed by a low pressure compressor 12 and subsequently by a high pressure compressor 17. Compressors 12 and 17 are of the axia1 type which produce an increase in pressure and conse-quently, an increase in the density of the incoming air proportional to their speedO The denser air moves to a fuel supply section where a fuel ring 16 produces a combustible $uel/air mixture by spraying fuel into the incoming stream~ The fuel/air mixture is thereafter ignited and combusted in burners 18 to produce a high velocity exhaust gas. The exhaust gas, as is conventional in this type of reaction engine, exits through a nozzle 20 to produce a forward thrust. A portion of the energy contained Tn the exhaust gas is expended in rotating a low pres-sure turhine 14 and a high pressure turbtne 19 which are physically attached to the respective compressors by concentrically disposed rotor shafts. Each compressor and turhine combination form a spool which rotates independently from the other. The engine described is a con-ventional two spool turboJet engine in which the second or low pres-sure spool has been added to increase the ef~iciency of the engine.
Generally, the control of the turbine speed and thus, output power or thrust of the en~ine is regulated by the amount of fuel flow-ing through the fuel ring 16. This fuel flow Wf is modulated by the fuel control to provide more or less energy and thus, reaction power from the engine. Increasing the fuel flow increases the engine spçed and output power while decreasing the fuel flow decreases engine speed and output power, The fuel control system 11 is provided for assuring that control of the engine speed is maintained during steady state opera-tion, accelerations, and decelerations~ The fuel control illustrated is based on an input to terminal 24 which is the desired or scheduled rotational speed Ngs, of one of the compressor and turbine combinations ~7~73~
of the engine 10. The signal Ngs can, for example, be generated as the output from a schedule based on the power lever angle of the particular device the engine is used on. In its simplest form the Ngs signal can be generated from a potentiometer that is ganged to the power lever.
When referring to the engine speed, either the high pressure compressor speed Ng or the low pressure compressor speed Nl is meant because either can be used to control the engine. In this particular embodiment the controlling parameter of the fuel control will be the high spool speed Ng. The high spool speed is the controlling parameter in this type of engine since it is the more critical to the operation.
~y scheduling the commanded speed Ngs to terminal 24 any desired engine power may be obtained since the fuel control varies the power of the gas turbine engine by regulating engine speed. Thus, the control system operates to govern the engine speed as a function of the scheduled speed Ngs and accelerates or decelerates the engine in a manner to reach a new scheduled speed when the pilot or another system component changes the scheduled engine speed. Otherwise, for changes in engine environment, such as altitude, ternperature, pressure, or other factors, the governor acts to maintain the scheduled speed Ngs.
To control acceleration or deceleration, the controller forms an acceleration term which is the difference between the scheduled speed Ngs and a demanded speed term Nd. This function is performed by feeding the two parameters to a first summing junction 26 and generating a differ-ence signal of sign and magnitude equlvalent to the difference there-between. The acceleration term or difference signal is integrated in an integrator 32 to become the demanded speed signal Nd. This portlon of the control forms an integral loop providing a datum which is the demanded speed signal Nd.
The acceleration term is conventionally limited by a compara-tor circuit 28. The comparator 28 compares the acceleration term to an upper limit Al and a lower limit Dl and pases the acceleration term unmodified if between the two values~ However, if the demanded accelera-tion is greater than that defined by the current operating conditions, the acceleration term will be limited at the upper acceleration limit Al, while if the demanded deceleration is greater than that defined by the 7~73~1 `
7 ~
current operating conditions it will be limited at the lower deceleration limit Dl. An aceleration/deceleration limiter circuit 30 is used to generate the variable limits and to modify the limits as a function of the operating parameters of the engine as is conventional in the art. An acceleration limit schedule is stored in a memory device of the circuit where the schedule is substantially equivalent to the surge line of the particular engine expressed as a function of the engine operating para-meters chosen. The surge line can be expressed as any of a number of different parameters indicating steady state values of fuel/air ratio, or acceleration ratios.
Thereafter, the demanded speed signal Nd is differenced with the high spool speed of the turbine Ng in a second summing circuit 34.
The error signal E generated by this difference is then used to propor-tionally control the fuel flow to the engine 10 in a direction tending to null the error between the scheduled speed signal Ngs and the demanded speed signal Nd. In this manner thts portion of the controller forms a proportional control loop slaved to the datum Nd. For changes in the datum Nd the actual engine speed Ng will change to where there is a steady state error E supporting that particular demanded speed.
The proportional control is developed by a hydromechanical untt (HMU~ 36 which receives as an Tnput the error signal E. The HMU 36 multiplies the error signal E by a predetermtned gain to provide a fuel flow Wf which is proportional to the error slgnal. In its simplest form, if the error signal E is electrical, the HMU could be implemented as a proportional solenoid valve which changes position with respect to a control voltage representative of the error signal to regulate fuel flow from a pressurized fuel source. The HMU 36 wouldj therefore, also consist of a pressure regulator and a pressurized source of fuel (not shown~ as is conventionally known in the art.
Additionally, the HMU 36 can include means for meshanically multiplying the error signal E by an input representative of the outpu~
of the compressor stages, compressor pressure Pc. In combination with the proportional solenoid, the multiplying means could comprise a multiplying piston coupled to the fuel valve which transduces the compressor pressure Pc into a force to assist the solenoid. Thus, ~1~273 the error signal E at all engine operating points is proportional to Wf/Pc or the fuel/air ratio of the engine. Because of this mechanical multiplication, the error signal E can be thought of as being generated as a function of the fuel/air ratio of the engine. Normally, this would make the scheduling of the speed signal Ngs much easier to imple-ment.
The governing response of the fuel control system will now be more fully explained with respect to Figure 3. Illustrated is a curve that represents the steady state values of Nd for all values of steady state speed Ng. The curve Nd is linear and monotonic with a predetermined slope. A slope on the Nd curve generates greater error signal values E for increases in speed. For example, E2 is greater than El where N2 is greater than Nl. This is to be expected where the s~eady state fuel flow must increase for adjustments in engine speed to maintain equilibrium.
To accelerate the engine from Nl, the signal Nd is controlled by the integrator to increase above its steady state value thereby increasing the error signal above its steady state value. The amount of overfueling produced by the increase in the error signal is propor-tional to the acceleration of the engine As the engine increasesspeed due to the overfuellng, the Ng signal will increase to where it reduces the error signal to a new equilibrium value E2, Conversely, reducing the error signal below its steady state value by decreasing the Nd signal will cause an underfueling and a consequent deceleration of the engine from N2. The engine will decelerate until the Ng signal decreases to where the error signal El is again in equilibrium between the two values.
It is evident that the Ng signal, and thus the actual sp~ed of the turbine, follows the Nd signal proportionally wlth a lag based on the inertial response of the engine to the increases and decreases in fuel flow. Therefore, the actual speed of the rotor can be controlled ~y using the Nd signal as an anticipatory or implied actual speed signal to govern engine performance.
For an appreciation of the improvement in system perform-ance by using the Nd signal as an anticipatory feedback, Figure 4 will now be explained in conjunction with certain aspects of Figure 3.
7~7~
Figure 4 illustrates the time traces of a transient response to an acceleration request for a closed loop system based on acceleration where in one instance the acceleration term is based on actual rotor speed and in the other instance the acceleration term is based on demanded rotor speed. In other words, the controller illustrated in Figure 1 will be compared to the prior art slave datum controller.
The two traces indicate the relationship of Nd and Ng on a time base for a transient ~hat produces an acceleration to move the actual speed of the turbine from Ng~l) to Ng~2). Of course, this will require a change in demanded speed from Nd~l) to Nd(2~. This change in turbine speed corresponds to a change in steady state conditions from Nl to N2 in Figure 3 and thus El and E2 are identical For both Figures.
Now assume the scheduled rotor speed Ngs is changed in a step-wise fashion from a value corresponding to Ng(l) to a value corresponding to Ng(2~. The integrator will begin integrating the difference between Ngs and Nd in the present system and the difference between Ngs and Ng in the prior art system. The systems will continue to change fuel flow until the differences are integrated to zero~ i.e., until Ngs = Nd in the present system and Ngs - Ng in the prior art system.
It is reasonable to predict a large scale transient represent-ing a relatively large difference signal will transfer both systems to a limittng acceleration value. While on the acceleration limit the Nd signal will increase according to a predetermined slope set by the limit and the Ng signal will follow it by its inertial time lag. It is assumed that the operating conditions input to the limiter circuit 32 remain constant so that the acceleration limit does not vary during this time.
The system will continue to accelerate at the limit value until the difference signal or acceleration term is reduced below the limit value. At that time, point A in the drawing, the demanded speed signal Nd will begin to slow its increase as the integrator begins integrating smaller difference signals. When the demanded speed signal reaches point B, it equals the Ngs signal and the integrator thus holds it constant since the difference signal is now zero, This is prior to the actual speed signal Ng reaching its steady state value. The system subsequently comes into equilibrium when the actual speed signal Ng ~L7~73~ :
proportionally following the demanded spaed signal Nd reaches point C
an inertial time delay later.
This system thus provides a rapid but smooth trans;tion for a change in speed from Ng(l~ to Ng(2~. It is noted that the actual speed signal Ng slows hefore it reaches Ng(2) so that no overshoot is produced. When transitioning from the acceleration limit the system response is similar for large and small transient conditions because the Nd signal will always perform in an identical manner near the equilibrium value.
The normal slave datum system, however, when point A is reached continues to accelerate to point D where the difference between the actual speed signal Ng and the scheduled speed signal Ngs is less than the acceleration limit. Since there is a lag in the response of the Ng signal to the Nd signal, the system does not recognize until this point that it should lower the acceleration to a smaller amount.
Consequently, the Nd signal does not stop demanding an increase in acce1eration until at point E one time delay later than the present system. By that time, the Nd signal which has demanded more over-fueling than necessary pulls the Ng signal past the scheduled Ngs signal and requires a deceleration cycle to be generated unttl the system finally settles at point F.
Depending upon the inertTal lag of the engine and the pro-portional time constant (gain) of the system this response may not only overshoot the steady state value, but also be oscillatory. The only way to control this oscillatory transient response is ~o decrease the proportional gain. This, however, produces a more sluggish system that does not respond optimally to transients. Therefore, it has been shown that the use of the demanded speed signal to generate an implied acceleration term may be used to increase response to transients while maintaining loop stability in the gas turbine engine fuel controller.
The fuel control system that has just been described was shown in Figure 1 implemented as specifically applicable to a turbo-jet engine. This fuel control system is just as advantageous in the control of a free turbine engine as will now be more fully pointed out with reference to Figure 2. In that Figure, a fuel control system 104 273~i is shown as receiving a scheduled speed signal Ngs, an actual speed sTgnal Ng, and a pressure signal Pc, and generating a ~uel flow sig-nal Wf therefrom. The components comprising fuel control system 104 are identical to that described for the system 11 of Figure 1.
However, in this implemen~ation the signal ~f controls a gas generator 106 for a free turbine engine. The gas generator 106 contains similar components to the turbojet engine including a compressor7 a fuel ring, a burner section, and a turbine. However, instead of using the high velocity gas in a reaction generator, the high energy gas is used to turn a power turbine 108 of the free turbine engine. It will be appreciated that the power turbine can be connected to any type of load normally powered by a free turbine engine.
When the invention is utilized in this manner the scheduled speed signal Ngs, which is a measure of the output power of the gas generator, can be scheduled by a scheduling circuit 102. Advantageous circuits for generating the scheduled speed o~ the gas generator when the invention controls a free turbine engine can take many forms.
In one specific implementation, the scheduling ctrcuit 102 can have an input signal Nf from t~e power turbine which indicates the rotational speed thereof. The implementation can also have a variable reference signal Nr which is representative of the desired speed of the power turbine 108 input to the schedultg circuit 102. A
proportional control loop can 6e formed by dîfferencing the signals Nr and Nf generating the signal Ngs proport~onally to the error produced therefrom. Thus, the scheduling circuit can be a free turbine speed governor providing a reset signal to a gas generator speed governor, fuel control system 104.
In another more specific implementation, the power turbine 108 can be connected through suitable gearing or a transmission to a heli-copter rotor 110. The scheduling circuit 102 in this implementation would comprise a power turbine speed governor which maintains the free turbine speed against the reference signal Nr as previously described.
~owever, since it is desired that the rotor 110 maintain a substantially constant speed, in this instance the Nr signal will be a substantially constant valueO An additional input to govern the output power of the ~;~7~7~ :
gas generator can be a signal CLP indicative of the collective pitch angle of the rotor. This signal will therefore be representative of the desired output power of the engine as demanded by a pilot while maintaining rotor speed constant.
While a detailed description of the preferred embodiments has been disclosed, it will be obvious to those skilled in the art that various modifications and changes may be made thereto without departing from the spirit and scope of the invention as is herein-after defined in the appended claims.
~273~
FUEL CONTROL SYSTEM FOR GAS TURBINE ENGINE
BACKGROUND OF THE INVENTION
-- . . ;
The invention relates to a fuel control system for a gas turbine engine.
For the improved fuel control of gas turbine engines it is known to use a closed loop govenor based on the speed of the turbine rotor. Proportionai governing is introduced in these controls by differencing the actual speed of the turbine and a demandec' or scheduled speed. The error signal developed by differencing these parameters is multiplied b~/ a gain and thereafter used to control fuel flow to change turbine speed in a direction tending to null the error. A pure pro-portional type of cortrol, however~ produces a steady state or perma-nent error between the actual and desTred speeds which increases with turbine speedO This permanent error or "droop" is indicative of the steady state fuel flow at the quiescent operating points of the control-ler. Although the error is inversely proportional to the gain of the proportional loop, it can only be truly nulled with a loop of infinite gain. However, Increasing the gain of a proportional loop to where the "droop" error becomes acceptable can result in instability and loss of control of speed for the gas turbine engine.
To overcome this~ problem, isochronous governors for gas turbine engines have been proposed where the permanent error signal is integrated to zero without producing instability in the proportional loop. The deviation or "droop" error between the scheduled and actual speed signals can then be effectively nulled in this manner. Implemen-tation of tha techni~ue comprises an integral loop in parallel with he proportional loop. The integral loop includes an integrator for ~enerating the time integral of the difference between the scheduled speed and the actual speed. In this manner the integrator output yields a datum to which the proportional loop can be slaved by differencing ;t with the actual speed. Thus, it is the difference between the datum and the actual speed which generates the permanent error signal and steady state fuel control value. In this type o~ system the actual speed becomes slaved to the datum and changes proportionally according to the loop gain with changes in the datum.
~ ~7273~
When the controlling parameter of a slave-datum system is a scheduled speed, the deviation input to the integral loop is an acceleration term since the time integral of acceleration is velocity.
The system is closed loop on acceleration where the magnitude of the difference between the scheduled speed and the actual speed is indica-tive of a scheduled acceleration. The actual acceleration of the engine is fed bac~ through the changes in the actual speed signal to close the loop. It is evident that acceleration limiting of the gas turbine engine can be accomplished in a facile manner by regulating the accelera-tion term before it is input to the integrator~
Although this type of isochronous system provides the advan-tages of being able to null the droop of the governor and limit accelera-tion it does have one disadvantage. The system response to changing transient conditions is regulated by the response of the gas turbine engine to thcse conditions. Before the error caused by a change in operating condttions or a scheduled acceleration can be integrated to a nullity9 the engine must actually change speed an amount equal to the error. Therefore, changes in the datum to which the proportional loop is slaved are dependent upon the inertial lag of the engine. It would be highly desirable to remove this inertial time delay from the integral loop.
It has now been determined that the actual turbine speed will proportionally follow the demanded speecl sign31 or datum by the inertial lag of the engine and thus it can be assumed that the demanded speed signal contains useful information about the anticipated rate of change of actual turbine speed. This assumption is reliably based on the premise that the engine response characteristic is known and it will change speed as commanded by the demanded speed signal. An anticipatory estimate of that rate of change of actual speed can be obtained by comparing the scheduled speed signal to the demanded speed signal ~o generate the acceleration term, thus removing the engine lag from the integral loop calculation.
SUMMARY OF THE INVENTION
The invention is an improved closed loop fuel control system for gas turbine engines. The invention comprises a proportional loop ~7~7~
which differences a demanded speed signal (datum) and an actual speed signa1 and an integra1 loop which integrates an acceleration term to yield the demanded speed signal. The proportional loop is slaved to the demanded speed signal to generate an error signal which controls fuel flow of the engine in a direction tending to null the acceleration term. The acceleration term is integrated to zero for steady state con-ditions by the integral loop. The acceleration term instead of being formed in the normal manner is formed by differencing a scheduled speed signal and the demanded speed signal. This system is closed loop on acceleration because the rate of change of the demar,ded speed signal is utilized as an anticipatory estimate of actual acceleration. Therefore, the acceleration term will lead the actual acce1eration as the demanded speed signa1 1eads the actua1 speed signa1.
By comparing the scheduled speed signa1 to the demanded speed signa1 a higher proportiona1 loop gain may be used in the gas turbine engine fue1 contro1 without producing instabi1ity. This is because the demanded signa1 anticipates the steady state va1ue to which the engine is acce1erating and does not overshoot this va1ue.
Therefore, the actua1 engine speed which proportionally fol10ws the demanded speed signa1 by the inertia1 la~ of the engine will also begin dece1erating before it reaches Its steady state va1ue. The gain of the proportiona1 100p can be increased because the datum is reduced automatica11y upon nearing the steady state value and will not cause the slaved actual speed to overshoot the value.
In a preferred embodiment the improved fuel control is used in conjunction with a hydromechanical unit that regulates fuel flow proportiona11y in dependence on the error signa1. The error signa1 is generated by the contro1 as a function of Wf~Pc where Wf is fuel f10w and Pc is compressor pressure. The hydromechanical unit receives `the error signal in term of the fuel~air ratio and mechanical1y multi-p1ies the signal by a parameter based on compressor pressure to modu-1ate fue1 flow to the engine.
In the preferred embodiment the acce1eration term is also 1imited prior to integration by an acce1eration and dece1eration 1imit based upon the surge 1ine of the particu1ar engineO This configuration ~L72~30 is particularly advantageous in that for transient response the integrator can be set on limit until the acceleration term is reduced to below the limit, whereafter the demanded speed signal can smoothly approach the desired steady state value in a controlled manner. In spite of wherever the original transient begins or its size, this fea~ure produces the same final behavior from the engine. Thus, the fina1 transition from the acceleration schedule to a steady state value can be tailored to bring the system into equilibrium in an optimum amount of time without overshoot.
Another feature of the invention provides that the scheduled speed term is generated by a speed scheduling circuit as a function of the demanded output power of the gas turbine engine. In one pre-ferred implementation the speed scheduling circuit is a frea turbine speed governor for a shaft turbine engine. In a second preferred implementation the speed scheduling circuit governs the substantially constant rotor speed of a helicopter rotor with respect to a reference speed. The output power of the shaft turbine engine in the second implementation is controlled by the operator as a function of the collective pitch angle of the rotor.
These and other objects, features, and aspects of the inven-tion will be better understood and more clearly explained if a reading of the detailed disclosure is undertaken in conjunction with the appended drawings, wherein:
BRIEF DESCRIPTION OF THE DRAWINGS
Figure 1 is a system block diagrammatic vtew of a fuel control system for a turhojet engine which is constructed in accordance with the teachings of the invention; and Figure 2 is a block diagrammatic view of the fuel control system illustrated in Figure 1 in an embodiment suitable for a shaft 3 turbine engine;
Figure 3 is a graphical representation of the steady state values of the demanded speed signal and the actual speed signal as a function of quiescent turbine speed for the fuel control system illus-trated in Figure l; and ~7~730 Figure 4 is a time trace comparison of the demaned speed signal and actual speed signal for the system disclosed in Figure 1 and a prior art system.
DETAILED DESCRIPTION OF THE PREFER~ED EMBODIMENTS
.. . ..
With reference to Figure 1 there is shown a fuel control system 11 for a turbojet engine which is generally designated 10. The engine 10 includes an intake duct 2~ through which air enters. After entering the engine7 the air is compressed by a low pressure compressor 12 and subsequently by a high pressure compressor 17. Compressors 12 and 17 are of the axia1 type which produce an increase in pressure and conse-quently, an increase in the density of the incoming air proportional to their speedO The denser air moves to a fuel supply section where a fuel ring 16 produces a combustible $uel/air mixture by spraying fuel into the incoming stream~ The fuel/air mixture is thereafter ignited and combusted in burners 18 to produce a high velocity exhaust gas. The exhaust gas, as is conventional in this type of reaction engine, exits through a nozzle 20 to produce a forward thrust. A portion of the energy contained Tn the exhaust gas is expended in rotating a low pres-sure turhine 14 and a high pressure turbtne 19 which are physically attached to the respective compressors by concentrically disposed rotor shafts. Each compressor and turhine combination form a spool which rotates independently from the other. The engine described is a con-ventional two spool turboJet engine in which the second or low pres-sure spool has been added to increase the ef~iciency of the engine.
Generally, the control of the turbine speed and thus, output power or thrust of the en~ine is regulated by the amount of fuel flow-ing through the fuel ring 16. This fuel flow Wf is modulated by the fuel control to provide more or less energy and thus, reaction power from the engine. Increasing the fuel flow increases the engine spçed and output power while decreasing the fuel flow decreases engine speed and output power, The fuel control system 11 is provided for assuring that control of the engine speed is maintained during steady state opera-tion, accelerations, and decelerations~ The fuel control illustrated is based on an input to terminal 24 which is the desired or scheduled rotational speed Ngs, of one of the compressor and turbine combinations ~7~73~
of the engine 10. The signal Ngs can, for example, be generated as the output from a schedule based on the power lever angle of the particular device the engine is used on. In its simplest form the Ngs signal can be generated from a potentiometer that is ganged to the power lever.
When referring to the engine speed, either the high pressure compressor speed Ng or the low pressure compressor speed Nl is meant because either can be used to control the engine. In this particular embodiment the controlling parameter of the fuel control will be the high spool speed Ng. The high spool speed is the controlling parameter in this type of engine since it is the more critical to the operation.
~y scheduling the commanded speed Ngs to terminal 24 any desired engine power may be obtained since the fuel control varies the power of the gas turbine engine by regulating engine speed. Thus, the control system operates to govern the engine speed as a function of the scheduled speed Ngs and accelerates or decelerates the engine in a manner to reach a new scheduled speed when the pilot or another system component changes the scheduled engine speed. Otherwise, for changes in engine environment, such as altitude, ternperature, pressure, or other factors, the governor acts to maintain the scheduled speed Ngs.
To control acceleration or deceleration, the controller forms an acceleration term which is the difference between the scheduled speed Ngs and a demanded speed term Nd. This function is performed by feeding the two parameters to a first summing junction 26 and generating a differ-ence signal of sign and magnitude equlvalent to the difference there-between. The acceleration term or difference signal is integrated in an integrator 32 to become the demanded speed signal Nd. This portlon of the control forms an integral loop providing a datum which is the demanded speed signal Nd.
The acceleration term is conventionally limited by a compara-tor circuit 28. The comparator 28 compares the acceleration term to an upper limit Al and a lower limit Dl and pases the acceleration term unmodified if between the two values~ However, if the demanded accelera-tion is greater than that defined by the current operating conditions, the acceleration term will be limited at the upper acceleration limit Al, while if the demanded deceleration is greater than that defined by the 7~73~1 `
7 ~
current operating conditions it will be limited at the lower deceleration limit Dl. An aceleration/deceleration limiter circuit 30 is used to generate the variable limits and to modify the limits as a function of the operating parameters of the engine as is conventional in the art. An acceleration limit schedule is stored in a memory device of the circuit where the schedule is substantially equivalent to the surge line of the particular engine expressed as a function of the engine operating para-meters chosen. The surge line can be expressed as any of a number of different parameters indicating steady state values of fuel/air ratio, or acceleration ratios.
Thereafter, the demanded speed signal Nd is differenced with the high spool speed of the turbine Ng in a second summing circuit 34.
The error signal E generated by this difference is then used to propor-tionally control the fuel flow to the engine 10 in a direction tending to null the error between the scheduled speed signal Ngs and the demanded speed signal Nd. In this manner thts portion of the controller forms a proportional control loop slaved to the datum Nd. For changes in the datum Nd the actual engine speed Ng will change to where there is a steady state error E supporting that particular demanded speed.
The proportional control is developed by a hydromechanical untt (HMU~ 36 which receives as an Tnput the error signal E. The HMU 36 multiplies the error signal E by a predetermtned gain to provide a fuel flow Wf which is proportional to the error slgnal. In its simplest form, if the error signal E is electrical, the HMU could be implemented as a proportional solenoid valve which changes position with respect to a control voltage representative of the error signal to regulate fuel flow from a pressurized fuel source. The HMU 36 wouldj therefore, also consist of a pressure regulator and a pressurized source of fuel (not shown~ as is conventionally known in the art.
Additionally, the HMU 36 can include means for meshanically multiplying the error signal E by an input representative of the outpu~
of the compressor stages, compressor pressure Pc. In combination with the proportional solenoid, the multiplying means could comprise a multiplying piston coupled to the fuel valve which transduces the compressor pressure Pc into a force to assist the solenoid. Thus, ~1~273 the error signal E at all engine operating points is proportional to Wf/Pc or the fuel/air ratio of the engine. Because of this mechanical multiplication, the error signal E can be thought of as being generated as a function of the fuel/air ratio of the engine. Normally, this would make the scheduling of the speed signal Ngs much easier to imple-ment.
The governing response of the fuel control system will now be more fully explained with respect to Figure 3. Illustrated is a curve that represents the steady state values of Nd for all values of steady state speed Ng. The curve Nd is linear and monotonic with a predetermined slope. A slope on the Nd curve generates greater error signal values E for increases in speed. For example, E2 is greater than El where N2 is greater than Nl. This is to be expected where the s~eady state fuel flow must increase for adjustments in engine speed to maintain equilibrium.
To accelerate the engine from Nl, the signal Nd is controlled by the integrator to increase above its steady state value thereby increasing the error signal above its steady state value. The amount of overfueling produced by the increase in the error signal is propor-tional to the acceleration of the engine As the engine increasesspeed due to the overfuellng, the Ng signal will increase to where it reduces the error signal to a new equilibrium value E2, Conversely, reducing the error signal below its steady state value by decreasing the Nd signal will cause an underfueling and a consequent deceleration of the engine from N2. The engine will decelerate until the Ng signal decreases to where the error signal El is again in equilibrium between the two values.
It is evident that the Ng signal, and thus the actual sp~ed of the turbine, follows the Nd signal proportionally wlth a lag based on the inertial response of the engine to the increases and decreases in fuel flow. Therefore, the actual speed of the rotor can be controlled ~y using the Nd signal as an anticipatory or implied actual speed signal to govern engine performance.
For an appreciation of the improvement in system perform-ance by using the Nd signal as an anticipatory feedback, Figure 4 will now be explained in conjunction with certain aspects of Figure 3.
7~7~
Figure 4 illustrates the time traces of a transient response to an acceleration request for a closed loop system based on acceleration where in one instance the acceleration term is based on actual rotor speed and in the other instance the acceleration term is based on demanded rotor speed. In other words, the controller illustrated in Figure 1 will be compared to the prior art slave datum controller.
The two traces indicate the relationship of Nd and Ng on a time base for a transient ~hat produces an acceleration to move the actual speed of the turbine from Ng~l) to Ng~2). Of course, this will require a change in demanded speed from Nd~l) to Nd(2~. This change in turbine speed corresponds to a change in steady state conditions from Nl to N2 in Figure 3 and thus El and E2 are identical For both Figures.
Now assume the scheduled rotor speed Ngs is changed in a step-wise fashion from a value corresponding to Ng(l) to a value corresponding to Ng(2~. The integrator will begin integrating the difference between Ngs and Nd in the present system and the difference between Ngs and Ng in the prior art system. The systems will continue to change fuel flow until the differences are integrated to zero~ i.e., until Ngs = Nd in the present system and Ngs - Ng in the prior art system.
It is reasonable to predict a large scale transient represent-ing a relatively large difference signal will transfer both systems to a limittng acceleration value. While on the acceleration limit the Nd signal will increase according to a predetermined slope set by the limit and the Ng signal will follow it by its inertial time lag. It is assumed that the operating conditions input to the limiter circuit 32 remain constant so that the acceleration limit does not vary during this time.
The system will continue to accelerate at the limit value until the difference signal or acceleration term is reduced below the limit value. At that time, point A in the drawing, the demanded speed signal Nd will begin to slow its increase as the integrator begins integrating smaller difference signals. When the demanded speed signal reaches point B, it equals the Ngs signal and the integrator thus holds it constant since the difference signal is now zero, This is prior to the actual speed signal Ng reaching its steady state value. The system subsequently comes into equilibrium when the actual speed signal Ng ~L7~73~ :
proportionally following the demanded spaed signal Nd reaches point C
an inertial time delay later.
This system thus provides a rapid but smooth trans;tion for a change in speed from Ng(l~ to Ng(2~. It is noted that the actual speed signal Ng slows hefore it reaches Ng(2) so that no overshoot is produced. When transitioning from the acceleration limit the system response is similar for large and small transient conditions because the Nd signal will always perform in an identical manner near the equilibrium value.
The normal slave datum system, however, when point A is reached continues to accelerate to point D where the difference between the actual speed signal Ng and the scheduled speed signal Ngs is less than the acceleration limit. Since there is a lag in the response of the Ng signal to the Nd signal, the system does not recognize until this point that it should lower the acceleration to a smaller amount.
Consequently, the Nd signal does not stop demanding an increase in acce1eration until at point E one time delay later than the present system. By that time, the Nd signal which has demanded more over-fueling than necessary pulls the Ng signal past the scheduled Ngs signal and requires a deceleration cycle to be generated unttl the system finally settles at point F.
Depending upon the inertTal lag of the engine and the pro-portional time constant (gain) of the system this response may not only overshoot the steady state value, but also be oscillatory. The only way to control this oscillatory transient response is ~o decrease the proportional gain. This, however, produces a more sluggish system that does not respond optimally to transients. Therefore, it has been shown that the use of the demanded speed signal to generate an implied acceleration term may be used to increase response to transients while maintaining loop stability in the gas turbine engine fuel controller.
The fuel control system that has just been described was shown in Figure 1 implemented as specifically applicable to a turbo-jet engine. This fuel control system is just as advantageous in the control of a free turbine engine as will now be more fully pointed out with reference to Figure 2. In that Figure, a fuel control system 104 273~i is shown as receiving a scheduled speed signal Ngs, an actual speed sTgnal Ng, and a pressure signal Pc, and generating a ~uel flow sig-nal Wf therefrom. The components comprising fuel control system 104 are identical to that described for the system 11 of Figure 1.
However, in this implemen~ation the signal ~f controls a gas generator 106 for a free turbine engine. The gas generator 106 contains similar components to the turbojet engine including a compressor7 a fuel ring, a burner section, and a turbine. However, instead of using the high velocity gas in a reaction generator, the high energy gas is used to turn a power turbine 108 of the free turbine engine. It will be appreciated that the power turbine can be connected to any type of load normally powered by a free turbine engine.
When the invention is utilized in this manner the scheduled speed signal Ngs, which is a measure of the output power of the gas generator, can be scheduled by a scheduling circuit 102. Advantageous circuits for generating the scheduled speed o~ the gas generator when the invention controls a free turbine engine can take many forms.
In one specific implementation, the scheduling ctrcuit 102 can have an input signal Nf from t~e power turbine which indicates the rotational speed thereof. The implementation can also have a variable reference signal Nr which is representative of the desired speed of the power turbine 108 input to the schedultg circuit 102. A
proportional control loop can 6e formed by dîfferencing the signals Nr and Nf generating the signal Ngs proport~onally to the error produced therefrom. Thus, the scheduling circuit can be a free turbine speed governor providing a reset signal to a gas generator speed governor, fuel control system 104.
In another more specific implementation, the power turbine 108 can be connected through suitable gearing or a transmission to a heli-copter rotor 110. The scheduling circuit 102 in this implementation would comprise a power turbine speed governor which maintains the free turbine speed against the reference signal Nr as previously described.
~owever, since it is desired that the rotor 110 maintain a substantially constant speed, in this instance the Nr signal will be a substantially constant valueO An additional input to govern the output power of the ~;~7~7~ :
gas generator can be a signal CLP indicative of the collective pitch angle of the rotor. This signal will therefore be representative of the desired output power of the engine as demanded by a pilot while maintaining rotor speed constant.
While a detailed description of the preferred embodiments has been disclosed, it will be obvious to those skilled in the art that various modifications and changes may be made thereto without departing from the spirit and scope of the invention as is herein-after defined in the appended claims.
Claims (15)
1. A fuel control system for a gas tubine engine comprising:
means for generating an actual speed signal proportional to the actual turbine speed of the gas turbine engine;
means for generating a scheduled speed signal proportional to a desired engine output power;
means for generating a demanded speed signal proportional to the time integral of a difference signal;
means for generating said difference signal proportional to the difference between said scheduled speed signal and said demanded speed signal;
means for generating an error signal proportional to the difference between said demanded speed signal and said actual speed signal; and means for controlling the fuel flow to the gas turbine engine as a function of said error signal.
means for generating an actual speed signal proportional to the actual turbine speed of the gas turbine engine;
means for generating a scheduled speed signal proportional to a desired engine output power;
means for generating a demanded speed signal proportional to the time integral of a difference signal;
means for generating said difference signal proportional to the difference between said scheduled speed signal and said demanded speed signal;
means for generating an error signal proportional to the difference between said demanded speed signal and said actual speed signal; and means for controlling the fuel flow to the gas turbine engine as a function of said error signal.
2. A fuel control system as defined in claim 1, further including:
means for limiting said difference signal before integration by said demanded signal generating means between an upper limit and a lower limit.
means for limiting said difference signal before integration by said demanded signal generating means between an upper limit and a lower limit.
3. A fuel control system as defined in claim 2 wherein:
said upper limit is variable.
said upper limit is variable.
4. A fuel control system as defined in claim 3 wherein:
said lower limit is variable.
said lower limit is variable.
5. A fuel control system as defined in claim 1, wherein:
said error signal is generated as a function of Wf/Pc where Wf is the fuel flow delivered to the engine; and Pc is the compressor output pressure.
said error signal is generated as a function of Wf/Pc where Wf is the fuel flow delivered to the engine; and Pc is the compressor output pressure.
6. A fuel control system as defined in claim 5, wherein said fuel controlling means includes:
means for multiplying said error signal by said pressure signal to yield the fuel flow input to the engine.
means for multiplying said error signal by said pressure signal to yield the fuel flow input to the engine.
7. A fuel control system for a gas turbine engine of the free turbine type having a gas generator and a power turbine, characterized by:
means for generating an actual speed signal proportional to the actual turbine speed of the gas generator of the engine;
means for generating a scheduled speed signal proportional to the desired output power of the power turbine of the engine;
means for generating a demanded speed signal proportional to the time integral of a difference signal;
means for generating said difference signal proportional to the difference between said scheduled speed signal and said demanded speed signal;
means for generating an error signal proportional to the difference between said demanded speed signal and said actual speed signal; and means for controlling the fuel flow to the gas generator of the free turbine engine as a function of said error signal.
means for generating an actual speed signal proportional to the actual turbine speed of the gas generator of the engine;
means for generating a scheduled speed signal proportional to the desired output power of the power turbine of the engine;
means for generating a demanded speed signal proportional to the time integral of a difference signal;
means for generating said difference signal proportional to the difference between said scheduled speed signal and said demanded speed signal;
means for generating an error signal proportional to the difference between said demanded speed signal and said actual speed signal; and means for controlling the fuel flow to the gas generator of the free turbine engine as a function of said error signal.
8. A fuel control system as defined in claim 7, further including:
means for limiting said difference signal before integra-tion by said demanded signal generating means between an upper limit and a lower limit.
means for limiting said difference signal before integra-tion by said demanded signal generating means between an upper limit and a lower limit.
9. A fuel control system as defined in claim 8, wherein:
said upper limit is variable.
said upper limit is variable.
10. A fuel control system as defined in claim 9, wherein:
said lower limit is variable.
said lower limit is variable.
11. A fuel control system as defined in claim 10, wherein:
said error signal is generated as a function of Wf/Pc where Wf is the fuel flow delivered to the engine; and Pc is the compressor output pressure.
said error signal is generated as a function of Wf/Pc where Wf is the fuel flow delivered to the engine; and Pc is the compressor output pressure.
12. A fuel control system as defined in claim 11, wherein said fuel controlling means includes:
means for multiplying said error signal by said pressure signal to yield the fuel flow input to the engine.
means for multiplying said error signal by said pressure signal to yield the fuel flow input to the engine.
13. A fuel control system as defined in claim 7 wherein:
said power turbine operably powers a rotor of a helicopter.
said power turbine operably powers a rotor of a helicopter.
14. A fuel control system defined in claim 13, wherein:
said rotor speed is substantially constant and wherein said desired output power is demanded as a function of the pitch angle of the rotor.
said rotor speed is substantially constant and wherein said desired output power is demanded as a function of the pitch angle of the rotor.
15. A fuel control system for a gas turbine engine including an integrator for generating a demanded speed signal which is propor-tional to the time integral of an acceleration term; means for generating an error signal proportional to the difference between the demanded speed signal and an actual speed signal indicative of the actual speed of the engine; and means for controlling the fuel flow to the engine as a function of said error signal; said fuel control system characterized by:
means for generating said acceleration term as the difference between said demanded speed signal and a scheduled speed signal indica-tive of a desired engine output power.
means for generating said acceleration term as the difference between said demanded speed signal and a scheduled speed signal indica-tive of a desired engine output power.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US06/210,938 US4423592A (en) | 1980-11-28 | 1980-11-28 | Fuel control system for gas turbine engine |
| US210,938 | 1980-11-28 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1172730A true CA1172730A (en) | 1984-08-14 |
Family
ID=22784945
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000391111A Expired CA1172730A (en) | 1980-11-28 | 1981-11-27 | Fuel control system for gas turbine engine |
Country Status (4)
| Country | Link |
|---|---|
| US (1) | US4423592A (en) |
| EP (1) | EP0053547A1 (en) |
| BR (1) | BR8107625A (en) |
| CA (1) | CA1172730A (en) |
Families Citing this family (26)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4593523A (en) * | 1980-12-17 | 1986-06-10 | Allied Corporation | Method and apparatus for acceleration limiting a gas turbine engine |
| CA1177936A (en) * | 1981-04-24 | 1984-11-13 | Rowland M. Evans | Stress limiter apparatus for a gas turbine engine |
| US4532763A (en) * | 1983-01-28 | 1985-08-06 | General Electric Company | Isochronous gas turbine speed control |
| US4672806A (en) * | 1983-01-28 | 1987-06-16 | General Electric Company | Isochronous gas turbine speed control |
| US4594852A (en) * | 1984-10-30 | 1986-06-17 | General Electric Company | Aircraft engine control |
| US4993221A (en) * | 1988-12-21 | 1991-02-19 | General Electric Company | Gas turbine engine control system |
| US5134845A (en) * | 1990-07-23 | 1992-08-04 | United Technologies Corporation | Control for a gas turbine engine |
| US5361579A (en) * | 1993-06-01 | 1994-11-08 | Sundstrand Corporation | Turbojet fuel control system |
| US6811555B1 (en) * | 1996-09-16 | 2004-11-02 | Origin Medsystems, Inc. | Method and apparatus for performing anastomosis with eversion of tissue edges and joining of exposed intima of the everted tissue |
| US5868763A (en) | 1996-09-16 | 1999-02-09 | Guidant Corporation | Means and methods for performing an anastomosis |
| WO2001015607A1 (en) * | 1999-09-01 | 2001-03-08 | Origin Medsystems, Inc. | Method and apparatus for performing anastomosis |
| US6890338B1 (en) * | 2001-02-27 | 2005-05-10 | Origin Medsystems, Inc. | Method and apparatus for performing anastomosis using ring having tines with weak sections |
| US7431015B2 (en) * | 2006-01-31 | 2008-10-07 | Honeywell International Inc. | Fuel metering system proportional bypass valve error compensation system and method |
| US7712299B2 (en) * | 2006-09-05 | 2010-05-11 | Conocophillips Company | Anti-bogdown control system for turbine/compressor systems |
| GB0620310D0 (en) * | 2006-10-13 | 2006-11-22 | Rolls Royce Plc | Engine control |
| US8321119B2 (en) * | 2008-07-10 | 2012-11-27 | General Electric Company | Methods and systems to facilitate over-speed protection |
| US8224552B2 (en) * | 2008-07-10 | 2012-07-17 | General Electric Company | Methods and systems to facilitate over-speed protection |
| US20100005657A1 (en) * | 2008-07-10 | 2010-01-14 | Van Vactor David R | Methods and systems to facilitate over-speed protection |
| FR2976622B1 (en) * | 2011-06-16 | 2015-05-01 | Turbomeca | DOUBLE TURBOMOTEUR BODY ARCHITECTURE WITH HIGH PRESSURE COMPRESSOR CONNECTED TO LOW PRESSURE TURBINE |
| US9382847B2 (en) * | 2012-11-02 | 2016-07-05 | Pratt & Whitney Canada Corp. | Rotor resonance disturbance rejection controller |
| US9494085B2 (en) | 2015-01-19 | 2016-11-15 | United Technologies Corporation | System and method for load power management in a turboshaft gas turbine engine |
| US9909442B2 (en) * | 2015-07-02 | 2018-03-06 | General Electric Company | Method of controlling a position actuation system component for a gas turbine engine |
| US10584646B2 (en) * | 2016-07-29 | 2020-03-10 | Pratt & Whitney Canada Corp. | System and method for control of gas turbine engine |
| CN115075954B (en) * | 2022-05-23 | 2023-03-28 | 蓝箭航天空间科技股份有限公司 | Method and device for improving starting success rate of engine |
| CN117170225B (en) * | 2023-10-09 | 2024-04-09 | 清华大学 | Incremental controller for aircraft engines based on rotor acceleration feedback |
| CN118705062B (en) * | 2024-08-29 | 2024-12-06 | 中国航发湖南动力机械研究所 | Aviation turboshaft engine control method and device, electronic equipment and storage medium |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3006144A (en) | 1961-10-31 | arnett etal | ||
| US3152444A (en) | 1961-03-01 | 1964-10-13 | Bendix Corp | Fuel control apparatus for turbo-shaft engines |
| US3393691A (en) | 1965-12-13 | 1968-07-23 | Bendix Corp | Fuel control having proportional plus integral governor with variable proportional and integral gains |
| US3587229A (en) | 1969-03-17 | 1971-06-28 | Bendix Corp | Combustion engine fuel control having governor reset means |
| DE2162660C3 (en) | 1971-12-17 | 1974-09-12 | Motoren- Und Turbinen-Union Muenchen Gmbh, 8000 Muenchen | Control device for a gas turbine system, e.g. a motor vehicle or aircraft engine |
| US3832846A (en) | 1972-04-27 | 1974-09-03 | Woodward Governor Co | Speed governor with fuel rate control |
| GB1470757A (en) * | 1973-04-27 | 1977-04-21 | Lucas Industries Ltd | Fuel control systems for gas turbine engines |
| US3936226A (en) | 1974-06-07 | 1976-02-03 | United Technologies Corporation | Control system for variable pitch fan propulsor with reverse pitch |
| GB1519143A (en) | 1974-07-09 | 1978-07-26 | Lucas Industries Ltd | Electronic fuel control for a gas turbine engine |
| GB1520412A (en) * | 1974-07-09 | 1978-08-09 | Lucas Industries Ltd | Gas turbine engine fuel control systems |
| GB1520882A (en) | 1974-07-24 | 1978-08-09 | Lucas Industries Ltd | Electronic fuel control for a gas turbine engine |
| SE393160B (en) | 1975-08-29 | 1977-05-02 | Stal Laval Turbin Ab | CONTROL SYSTEM WITH EMERGENCY EQUIPMENT FOR GAS TURBINE SYSTEM |
| US4134257A (en) | 1976-11-03 | 1979-01-16 | The Garrett Corporation | Gas turbine fuel delivery and control system |
| GB1573095A (en) | 1976-11-26 | 1980-08-13 | Lucas Industries Ltd | Fuel control for a multi engine gas turbine installation |
| DE2702774C3 (en) | 1977-01-24 | 1980-12-18 | Bodenseewerk Geraetetechnik Gmbh, 7770 Ueberlingen | Device for speed control of turbo air jet engines |
| US4188781A (en) | 1978-04-25 | 1980-02-19 | General Electric Company | Non-linear dual mode regulator circuit |
-
1980
- 1980-11-28 US US06/210,938 patent/US4423592A/en not_active Expired - Lifetime
-
1981
- 1981-11-24 EP EP81401859A patent/EP0053547A1/en not_active Withdrawn
- 1981-11-24 BR BR8107625A patent/BR8107625A/en unknown
- 1981-11-27 CA CA000391111A patent/CA1172730A/en not_active Expired
Also Published As
| Publication number | Publication date |
|---|---|
| BR8107625A (en) | 1982-08-24 |
| EP0053547A1 (en) | 1982-06-09 |
| US4423592A (en) | 1984-01-03 |
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