IL91682A - Control system for gas turbine engines - Google Patents

Description

91682/2 ΤΛ roomie ·) _)η¾ m -i r nyn CONTROL SYSTEM FOR GAS TURBINE ENGINES This invention relates to gas turbine aircraft engines, more particularly to engine control systems.

Gas turbine engine control systems are well known in the art. These control systems are specially configured for use with a particular civilian or military aircraft application. For example, fighter aircraft must be capable of undergoing violent maneuvers which require changes in engine power with corresponding changes in engine thrust to severely accelerate or decelerate the aircraft. To perform these maneuvers, the pilot must execute sudden power lever movements usually referred to as "bodies", "chops" or "snaps". These power lever movements produce extreme engine speed, temperature and air flow excursions . Engine controllers for fighter aircraft must provide maximum engine response as quickly as possible.

An example of an engine used in fighter aircraft is the F100 engine manufactured by Pratt & Whitney Aircraft, a division of The United Technologies Corporation, the assignee of the present application. The F100 engine is a multiple spool axial flow turbine power plant having a fan jet engine configuration. The engine is characterized by a fan or low compressor coaxial with a high compressor rotor. Both the fan and high compressor have vanes whose angles are adjustable while the rotor blades are moving. The engine also has a variable area exhaust nozzle.

During operation, the rotor speeds of the fan and high pressure compressor rotors will vary from a high speed (intermediate or militar power) to a low speed (part power or idle power) . To accomplish the dramatic changes in engine power output, not only are the fan and high compressor vanes changing position with the speed of the respective rotor, the variable exhaust nozzle is also changing area, as the amount of fuel provided to the engine combustor varies from a high value to a low" value. For example, a "chop" in desired thrust would schedule the engine to reduce power from a military power condition, e.g. 12000 lbs. of thrust, to 4, 000 lbs of thrust by reducing the amount of fuel to the combuster of the engine and increasing the exhaust nozzle area. The fan and high compressor rotor speeds, turbine temperature and air flow will decrease in accordance with the corresponding thrust profile. Similarly, there is a significant increase in rotor speed, temperature and air flow when the power plant undergoes a transition from part power to military power.

With existing jet engine controls there is a "bootstrap" process which occurs when there is a request for more or less power. Known controllers initiate a request for more or less power by increasing or decreasing fuel flow in response to the change in throttle power lever angle (PLA) . The change in fuel flow produces changes in power via combustor exit conditions. Only in response to these changes are the engine spool speeds changed and, in turn, is the compression system geometry adjusted, such as the position of the fan variable vanes (FW) , high compressor variable vanes (HCW) and exhaust jet nozzle area (AJ) . In transient power situations the result is lessened engine flexibility and slower engine response. In addi-tion, systems which control the gas path variable engine parameters in response to changes in fuel flow are characterized by non-optimum engine performance, as evidenced by higher fuel consumption and potential compressor system instability.

As is further well known and historically, the control for the gas turbine engine has typically adjusted fuel flow in attempting to optimize operation of the engine. In a typical installation the fuel control monitors a plurality of engine operating parameters and processes these signals to produce outputs that would be indicative of a desired engine operation while assuring that the engine avoids surging, overheating and rich and lean flame out . To achieve this goal the computer portion of the control manifests a control logic that represents the opera- tion. of the engine and continuously schedules fuel flow to reflect the setting of the power lever.

In engines, particularly of the military variety, the control also, independently, monitors engine variables to schedule the variable geometry portions of the engine such as inlet guide vanes, exhaust nozzles and the like, to likewise attain optimum operation for any given operation within the engine's operating envelope.

Hence, it is apparent that a change in one control function would affect the condition of others so that there would be constant iterations of each of the controls to assure optimum operation of each. For example, a change in the exhaust nozzle area would typically change the pressure within the engine, which pressure would be monitored by the fuel control, which in turn would manifest a change in the fuel control to ultimately adjust fuel flow to reflect this change. In this process the scheduling of the fuel flow either to increase or decrease fuel will occur even prior to the time it takes the variable geometry of the engine to react. This "bootstrapping" effect has been addressed above and is further described later.

The control intends to avoid the problem of "bootstrapping" attains a faster thrust response and improved surge margin by synchronously scheduling fuel flow and the variable geometries of the engine in response to a single parameter which is a function of certain engine and aircraft operating variables. In order for this type of active geometry control system to be a useful system, it must be able to attain a high degree of repeatability in assuring that for any given setting the control will return to a given steady state operating point in the operating curve after any transition excursion. During a transient excursion, the control logic will assure that the point of operation is identical to the setting of the power lever which request is desired thrust even though the engine operating variable changes as a result of wear and tear of the engine, power extraction or compressor air bleed .

This invention contemplates utilizing a corrected speed (RPM) of the low pressure spool ( ^) of a twin spool engine as the primary control parameter. As is the case in many of the military engines, the low pressure spool is only aerodynamicaliy coupled to the high pressure spool. In order to attain the optimum engine operation from a performance standpoint, the corrected rotor speeds for the •i high and low pressure spool must be proportional to each other for every given steady-state engine condition.

In fighters and other military aircraft, it is extremely important that a demand by the pilot for a change iri thrust produced by the engine be as fast and as accurate as possible. The aircraft's ability to undergo the violent maneuvers anticipated when operating in the combat box, for example, in fighter aircraft, bears directly on the performance of that aircraft. When a demand for a thrust change is initiated, for example, when the pilot exercises a bodie maneuver, i.e., a quick demand for a drop in thrust (decel) followed by an ' immediate demand for an increase in thrust (accel) or vice versa, the engine should attain the demanded thrust levels by decelerating to the desired thrust level before accelerating to the desired thrust level in the quickest time possible. With heretofore known control logic, maneuvers, such as these bodies as well as chops, are influenced by the constraints owing to the high inertia of the rotating spool. Since a thrust change necessitates a decrease or increase in RPM of the high pressure compressor, this high inertia adversely affects the time responsiveness of . the engine.

Needless to say, it is also extremely important that the engine operates as efficiently as possible to achieve good TSFC (thrust specific fuel consumption) and stable engine operating conditions, namely, avoiding surge, engine flame out and overheating.

It has been found that by controlling both the fuel and the fan variable vanes as a function of corrected low pressure compressor speed, both steady state and transient operations are enhanced. In transients, this invention contemplates locking in a fixed corrected low compressor speed (N^) , setting a target for the desired thrust and zeroing in on this target by adjusting the area of the fan variable vanes. This logic allows the high pressure compressor spool to adjust speed to a value corresponding to the targeted thrust (Fn) . A proportional plus integral controller assures that the corrected engine airflow (Wa) is properly attained while N-^'ls held constant. Once the target is reached, the active low compressor controller (ALCC) automatically trims the low pressure compressor, spool speed and hence the low compressor speed and FCW to return to the operating line at a value that is equivalent to the desired speed ratio of the high and low pressure compressor to assure optimum engine performance.

It has further been found that by controlling both the fuel and high compressor variable vanes as a function of corrected low pressure compressor speed, we can enhance both steady state and transient operations. In transients, this invention contemplates locking in a fixed corrected high compressor speed ( 2) , setting a target for the desired thrust and zeroing in on this target by adjusting the angle of the high compressor variable vanes. This logic allows the low pressure compressor spool to adjust speed to a value corresponding to the targeted thrust (Fn) . A proportional plus integral controller assures that the speed is properly attained while is held constant. Once the target is reached, the active high compressor controller (AHCC) automatically trims the high pressure compressor spool speed and HCVV to return to the operating line at a value that is equivalent to the desired speed ratio of the high and low pressure compressor to assure optimum engine performance.

It has further been found that by controlling both the fuel and high compressor variable vanes as a function of corrected low pressure compressor speed, we can enhance both steady state and transient operations. In transients, this invention contemplates locking in a fixed corrected high compressor speed (N2) f setting a target for the desired thrust and zeroing in on this target by adjusting the angle of the high compressor variable vanes. This logic allows the low pressure compressor spool to adjust speed to a value corresponding to the targeted thrust (Fn) . A proportional plus integral controller assures that the -j speed is properly attained while 2 is held constant. Once the target is reached, the active high compressor controller (AHCC) automatically trims the high pressure compressor spool speed and HCW to return to the operating line at a value that is equivalent to the desired speed ratio of the high and low pressure compressor to assure optimum engine performance.

An object of the present invention is to provide an active geometry control system for gas turbine engines which uses a single parameter to set overall engine thrust as a function of total engine air flow and pressure ratio.

Another object of the present invention is to provide an active geometry control system for gas turbine engines having synchronous scheduling of combustor fuel flow with gas path variable parameters by means of variable compression system geometry parameters.

Another object of the present invention is to provide an active geometry control system that modulates compressor system air flow while minimizing excursions of engine rotor speeds.

According to the present invention, a system for controlling power from an aircraft engine having gas path variable engine components, variable area exhaust nozzle, and a burner for generating hot exhaust gases, the system includes a means for receiving aircraft parameter , signals, a means for receiving, engine parameter signals, including signals indicative of gas path variable engine components exhaust area, and fuel flow to the burner. A controller is included which receives the parameter signals and provides, in response through signals indicative of selected engine power level controls, signals to the engine that select the magnitude of the burner fuel flow and the exhaust nozzle area synchronously with the magnitude of the gas path variable engine components.

A further object of this invention is to provide improved control logic for a two spool fan jet engine powering aircraft having variable area vanes on the fan or on the high spool that assures fast thrust response responsive to power lever input while assuring optimum TSFC when the en-gine operates in the quiescent state.

A feature of this invention is control logic of a two spool gas turbine engine that controls fuel flow and the fan variable vanes as a function of a low pressure compressor speed parameter, which parameter is a function of Mach No., engine inlet total pressure and temperature and the positon of the power lever. Transient conditions (accel and decel) are attained by locking in , targeting Wa to attain the thrust level dictated by the demand of the power lever, and steady state conditions are attained by synchronously 'trim-ming the speed of the high pressure compressor and the position of the fan variable vanes until designed low-to-high pressure compressor speed condition is reached.

Another feature of this invention is to utilize a timer that may be reset as a function of aircraft or engine opera-ting variables to trim the control to a steady state condition once a thrust target has been achieved.

Another feature of this invention is to provide in control logic as described means for locking in the low pressure compressor speed as a function of the position of the power lever so that the lock condition will not deviate in deference to external influences, such as, wear and tear of the engine, and power extraction for aircraft accessories and the like.

A further feature of this invention is control logic of a twin spool gas turbine engine that controls fuel flow and the high compressor variable vanes as a function of a low pressure compressor speed parameter, which parameter is a function of Mach No., engine inlet total pressure and temperature and the position of the power lever. Transient conditions (accel and decel) are attained by locking in the high compressor spool speed, targeting low. compressor spool speed to attain the thrust level dictated by the demand of the power lever, and steady state conditions are attained by synchronously trimming the speed of the high pressure compressor and the position of the high pressure compres-sor variable vanes until the designed low-to-high pressure compressor speed condition is reached.

Another feature of this invention is to utilize a timer that may be reset as a function of aircraft or engine operating variables to trim the control to a steady state condition some time after a thrust target has been achieved. Another feature of this invention is to provide in control logic as described means for locking in the high pressure compressor speed as a function of the position of the power lever so that the lock condition will not de-viate in deference to external influences, such as, wear and tear of the engine, and power extraction for aircraft accessories and the like.

The foregoing and other features and advantages of the present invention will become more apparent from the fol- lowing description and accompanying drawings, wherein; Fig. 1 is a simplified schematic illustration of a gas turbine engine employing an active geometry control system provided according to the present invention.

Fig. 2 is a detailed schematic illustration of a por- tion of the active geometry control system of Fig. 1.

Fig. 3 is a simplified block diagram of a second portion of the controller of Fig. 1; Fig. 4 is a simplified block diagram of a third portion of the controller of Fig. 1 ; Fig. 5 is a diagrammatic illustration showing the relationship between high compressor vane position angle as a function of a high compressor corrected speed; Fig. 6 is a diagrammatic illustration showing transient to steady state response of the control system of Fig. 1 ; • Fig 7. is a diagrammatic illustration relating engine net thrust to exhaust area for selected fan rotor speeds; Fig 8. is a graphical representation of the low spool operating line plotting the low compressor variable vane (LCW) to the low pressure compressor speed (Ν·|) corrected for its inlet conditions illustrating the features of this invention; Fig. 9 is a schematic and block diagram illustrating the overall relationship of the control to the gas turbine engine and illustrating the control logic of the low compressor variable vanes; Fig. 10 is a graphical representation of the high spool operating line plotting the high compressor variable vane (HCVV) to the high pressure compressor speed (^) corrected for its inlet conditions illustrating the features of this invention; and Fig. 11 is a schematic and block diagram illustrating the overall relationship of the control to the gas turbine engine illustrating the control logic of the high compressor variable vanes.

Referring now to Fig. 1 there is illustrated in schematic form a jet engine 10 which is of a conventional twin spool type having a. tow pressure spool 12 and a high pressure spool 17.

The_low pressure spool 12 consists of the combined fan/low pressure compressor driven by a low pressure turbine 16. The high pressure spool 17 consists of the high pressure compressor 18 driven by the high pressure turbine 20. The high pressure compressor 18 and high pressure turbi ne 20 are sometimes referred to as the gas generator or engine core.

A conventional burner 22 , disposed between the compressor exit and turbine inlet serves to heat and accelerate the engine's working medium in order to energize the gas sufficiently to power the turbines and generate thrust. The high pressure s.pool 12 and low pressure spool 17are not mechanically connected to each other but rotate independently. The engine also includes an augmentor 24 receiving discharged gas from the low turbine. The gas exits the engine via an exhaust nozzle 26. As is conventional, an actuator 28 is used to control the position of the exhaust nozzle and thereby vary the area (AJ) of the exhaust discharge opening .

Also illustrated in Fig. 1 is an active geometry control system 29, described hereinafter. As is conventional, nous with the control of variable gas path parameters.

Both the fan and high compressor have a plurality of ... vanes whose position are adjustable from closed to fully open. As is well known, these variable vanes can be adjusted according to preprogrammed _schedules which will optimize the power and response of the engine. The pilot controls engine power output by varying the position or angle of throttle lever 32. The angle of the throttle lever as well as the rate of change of throttle lever angle is determinative of the amount of power supplied by the engine. Signals indicative thereof are provided on lines 34 to the controller where the power lever angle and the rate of change of power lever angle can be determined. Signals are provided to the jet engine for the system 29 to control engine parameters as illustrated by lines 36,38 and 40.

Referring now to Fig. 2, there is a detailed schematic illustration of the active geometry control system of Fig. 1. Modern aircraft employ an electronic engine control which monitors aircraft parameters and which is preprogrammed with a plurality of parameter schedules for selecting fuel flow and other engine parameters in accordance therewith. Although it is preferable to embody the present in- vention in a digital electronic form, those skilled in the art will note that the present invention can alternatively be embodied in analog electronic, hydraulic or mechanical means with appropriate conventional changes in hardware and software .

In the best mode, embodiment of the control system is digital and includes such hardware and software as -is needed to accomplish the functions detailed herein. Parameter sensors, digital-to-analog and analog-to-digital converters and conventional computer means which may be required are not illustrated for the sake of clarity. Also in the control system of Fig.2, a plurality of function generators have been diagrammatically shown with unlabeled axes for purposes of clarity. As is conventional, the ordinate parameter for each function generator is labeled after its symbol, while the abscissa parameters precede it. Further, the several parameters illustrated would be recognized by those skilled in the art to correspond to well known parameters which may be corrected (C) for certain standard positions (2,2.5) within the engine. Therefore, 2C2.5 (meas ) represent the measured value of 2 spool speed corrected for position 2.5 in the. engine. A glossary of terms is provided below.

GLOSSARY OF TERMS AJ jet nozzle area AJT transient area adjustment PLA power lever angle PT2 signal indicative of engine inlet pressure TT2 engine inlet temperature EPR engine pressure ratio Nl 2 fan rotor corrected speed at station 2 DN^ differential fan rotor corrected speed (error D J-, ,2 differential exhaust nozzle "area PgMIN minimum burner pressure PB burner pressure DPB burner pressure error FW fan variable vane position HCW high compressor variable vane position -,C2R fan rotor corrected speed request HCWREF high compressor variable vane position (Ref ) DHCVV differential corrected high compressor variable vane position HCW final requested high compressor variable vane position (Ref) DN2C2 5 differential high compressor corrected speed (error) GVV gain ^2^2.5 derivative corrected high compressor speed WFgg fuel flow request WF/PB fuel flow ratio ^1 low compressor rotor speed 2 high compressor rotor speed 2 engine inlet station 2.5 fan discharge station 3 compressor discharge station N2C2 5 high compressor speed corrected at station 2.5 As detailed hereinafter, the active geometry control system provided according to the present invention is characterized by independently coordinated use of main combus-tor fuel flow with selected geometries of variable gas path elements to optimize engine performance. Consequently, thrust response. and compression system stability of turbo jet and turbo fan gas turbine engines are improved. As indicated hereinabove, known jet engine control systems would provide fuel to the combustor in response to changes in the power lever or throttle angle initiated by the pilot. The increased fuel would produce changes in the engine dynamics, requiring changes in fan or low compressor variable vane position and high compressor variable vane position in addition to the jet nozzle area. The process of increasing (or decreasing) fuel flow before rotor speeds and hence, any gas path elements can sufficiently react is known to those skilled in the art as "boot strapping". This process limits the rate at which power transients can be produced by the jet engine.

.However, with the active geometry control system provided according to the present invention, fan and high compressor variable vane position, jet nozzle area and fuel flow are all synchronously controlled in response to a power request. The system anticipates the response of the engine such that either or both compression system air flows are modulated with reduced excursion or spool speeds. Fig. 2 illustrates the portion of the control system which gene- rates control signals for both the actuators which control the high compressor variable vane position and the fuel flow in the combustor. Fig. 3 is seen to detail that portion of the control system which generates signals for adjusting the area of the exhaust gas nozzle. Fig. 4 schematically details that portion of the control system which generates fan variable vane position control signals.

In order to adjust the high compressor variable vane position, the control system must compute an engine air flow parameter value as well as determine whether the engine is · accelerating or decelerating, and the magnitude thereof.

Throttle or power lever 32 is the principal mechanism by which the aircraft pilot requests changes in engine power. Signals indicative of the power lever angle (PLA) are provided to engine airflow schedule mechanism 44 which also re- ceives a plurality of signals from the aircraft, including engine inlet pressure (PT2 ) , airplane speed (Mach number) and engine inlet temperature (TT2 ) . The scheduled value of engine airflow is received by function generator 46, which generates a corrected value of the spool speed (N.jC2R) .

Function generator 48 outputs a requested value of high compression variable vane position (HCVV(Ref )) . The reference values (HCVV(Ref)) is summed at junction 50 with a differential value thereof (DHCW) comprised of a corrected high compressor speed sensor (D 2C2. (base ) ) generated at junction 52 from a reference value of corrected N2 spool speed (N2C2 5(re^)) anc^ a value of 2C2 5 measured just as the new power level was requested (N2C2 ^(lock)) . This corrected high compressor speed sensor value is converted by a gain variable vane (GVV) generator 54 as a function of N^C2R« Acceleration/Decelaration selector 56 receives signals indicative of the power lever angle as well as the measured value of the corrected h"2 spool speed, and generates therefrom values of the N. and N2 corrected spool speeds at the initiation of the power request ( ->C2.5 lock and N-.C2 lock) .

The lock value of N-.C-, cis summed at junction 58 with the measured, corrected 2 spool speed to produce a differential 2C2.5 value (ΕΝ2θ2,5> · This signal is provided to proportional and integral trim circuitry 60 which generates a cor- rection signal for HCW in a. conventional manner. Also included is circuitry- for detecting very rapid changes in the measured corrected N2 spool speed. Derivative circuitry 62 receives the measured corrected 2 spool speed and outputs a derivative value (d 2C2 5) which is provided to propor- tional trim circuitry 64 whose output signal of HCW is configured in conventional, proportional control manner. The ■ output signals of both the proportional trim circuitry 64 and proportional and integral trim circuitry 60 are provided to junction 66, where the signals are summed producing a trim or correction signal for the high compressor variable vane position (HCW (Trim) ) . This signal is provided to junction 68 where it is summed with (HCW ref) signal producing a target value of the high compressor variable vane position (HCW (target) ) .

- The control system provided according to the present invention compares the target value of the high compressor variable vane position with minimum and maximum allowable values determined by function generators 70 and 72. The value of the target high compressor variable vane position signal is adjusted accordingly if the target value exceeds the upper or lower bounds. This is the adjusted or base value of the high compressor variable vane position signal (HCW) base .

The control system of Fig. 2 is also characterized by slew request circuitry, including a deceleration timer 74, which receives signals indicative of the airplane's altitude and speed (Mach number) as well as an enable signal from Accel/Decel selector circuitry 56 that indicates the elapsed time from the initiation of the power lever command signal. As detailed hereinafter, the control system of Fig. 2 will automatically reconfigure the engine parameters to be in a steady state mode if the deceleration timer circuitry determines that the power level angle has not changed in a preset period (e.g. 20 seconds) . nal and provides a signal corresponding to the high compressor variable vane position at steady state (HCWSS) .

This signal is summed with a high compressor variable vane position request signal (HCW(req) ) to generate a high compressor variable vane error signal (HCVV(err)) . In response to a slew request · signal , slew error circuitry 78 provides the HCWERR signal to junction 80 where it is summed with the HCW(base) signal. As detailed hereafter, the slew re- quest circuitry modifies the HCVV(req) signal in dependence on the time elapsed from the initiation of a change in power level angle signal. If the throttle has been moved within a selected time period (e.g. 20 seconds) the engine is considered in a transient condition and the HCW(base) signal is adjusted accordingly. At the termination of the time pe-• riod the HCW(req) signal is modified by slew error circuitry 78 to slew the high compressor variable vane position back to the scheduled steady state position provided by function generator 76 and thereby reduce 2 rotor speed.

Rate limit circuitry' 81 receives the corrected HCW(base) signal and compares the rate of change of that signal to preselected rate limits scheduled by function generator 82 from the plane's altitude, speed, engine inlet temperature and engine inlet pressure (PT2). The. controller outputs a HCW request signal to the high compressor vane actuators.

A measured corrected value of the N-| or 2 spool speed is received by fuel flow circuitry 84,86 or, which, in con- . junction with the magnitude of the power lever angle, schedules fuel flow-to-burner pressure ratio signal. Rate li- miting circuitry 88 limits the fuel flow-to-burner pressure ratio request signal in accordance with function generator 90. The parameters used by the rate limit circuitry 88 are similar to those used by rate limit circuitry 82 in determining the limit of the high compressor variable vane re-- quest signal. A signal indicative of burner pressure is provided to junction 92 along with the fuel flow to burner mechanisms in the combustor (22, Fig. 1) .

Referring now to Fig. 3, the control system of Fig. 1 includes circuitry for selecting the area of the exhaust jet nozzle. For nonafterburner operation, function generator as a function of power lever angle the airplane's altitude, speed, the engine inlet temperature and pressure. This signal is provided to a jet area mode and trim selector circuitry 96 which provides the AJ base in dependence on the engine transient or steady state mode. Circuitry 96 receives signals indicative of the steady state value of the nozzle area. A transient nozzle area adjustment signal (AJ.(trans)) is scheduled by function generator 98 as a function of a core engine parameter, such as Ρ^/Ρ2 as well as seme of the same parameters as used for the AJ(base) signal. The AJ (trans) signal is primarily responsible for holding the N-jC? signal to a high value during deceleration for a specified time as set by timer circuitry 74 in Fig.2. The requested exhaust nozzle area is further modified at circuitry 96 by proportional and integral trim circuitry 100 and proportional trim circuitry 102. The N-]C2(lock) signal set by the accelerator/decelerator selector circuitry 56 in Fig. 2 is compared to the iC2 Fig. 4 is a simplified block diagram of that portion of the control system of Fig. 1 which generates a fan va-riabie vane position request signal for the fan vane actuators. The function generator 106 receives signals indicative of the corrected measured N-j spool speed (N^ C2> nd schedules therefrom a base value of the fan variable vane (FVV base) position signal. Signals indicative of the exhaust nozzle area as well as the maximum allowed exhaust nozzle area are received at junction 108. Function generator 110 outputs a trim authority index (KA) which is slewed from a value of zero to one, the zero value corresponding to zero FVV position trim. The circuitry insures that both the exhaust area and fan variable position trims will not be fully active at the same time to avoid possible unstable interactions. A differential value of the corrected measured N-| spool speed DN1C2' that generated from the difference between the N-jC2 lock signal value and the measured corrected N1 spool speed (NiC2( eas)) is provided to a proportional and integral trim circuitry 112 which, when summed with the trim authority index signal and the fan variable vane ^base signal, comprises the fan variable vane position request signal.

The system is configured such that as the exhaust nozzle is slewed to the steady state position or as the exhaust nozzle closes during a reacceleration , the differential fan variable vane position signal value (DFVV) is automatically trimmed to zero siftce the trim authority value is slewed to zero as well. The synchronization of nozzle area together with fuel flow, fan variable vane position and high compressor variable vane position accomplished through the use of throttle input provides a unique method of controlling both fan and high compressor operating lines, and marks a point of departure of the present invention from the prior art. Extremely fast engine thrust transients are possible since fan or low com- pressor rotor speed is held high. The system described herein allows total engine power to be controlled by air flow via compression system geometry and by fuel flow. The improved thrust response and compression system stability of of turbo jet. engines and turbo fan gas turbine engines that is provided by the present invention can be seen by reference to Fig. 5.

Fig. 5 is a diagrammatic illustration showing the steady state relationship between high compressor variable vane position (HCVV) and high compressor corrected speed (N2C"2.5) detailed by curve 114. If the engine is operating at high power (Point 116) and receives a command to decelerate to a low power (Point 118) the system provided according to the present invention configures the engine so that the high compressor corrected speed essentially remains constant. Subsequent power request to accelerate or decelerate correspond to excursions along line 120. Consequently, rapid thrust response can be obtained from the engine without waiting for the speed of the high compressor fan to increase or decrease. This instantaneous response can also be obtained at other power levels, as indicated by points 122 and 124 which form an excursion line 126.

The system provided according to the present invention is also capable of monitoring the engine operating condition. If the engine is in a transient condition, i.e. the time elapsed from the previous power request is less than a preselected value, the system provides the optimum combination of transient rate and compression system aerodynamic stability. If, on the other hand, the engine is in a steady state condition, the system provides the optimum combination of engine performance such as specific fuel consumption and compression system aerodynamic stabi.i ity by automatically trimming from the transient mode high compressor variable vane position to the steady state high compressor variable vane position, as illustrated in Fig.6. Curve 128 illustrates the relationship between high compressor variable vane position (HCW) and high compressor corrected speed ^2C2.5^ ^n a manner similar to that of Fig. 5. For example, if the engine is operating at high power (Point 130) and receives a command to decelerate the system locks the value of high compressor corrected speed(N2C2.5 excursion line 131 at temporary set point 132. The temporary set point is also on curve 134 which corresponds to an operating line of constant power request (N-|C2 =constant) . ■ If the subsequent power request has not been recieved within the selcted time period, the system trims the high compressor corrected speed and high compressor variable vane position to a value thereof corresponding to optimum engine performance (Point 136) .

· An advantage of synchronous exhaust area scheduling provided according to the present invention compared with prior art control systems is depicted in Fig. 7. Curves 138, 140,142 and 144 show net propulsive force as a function of exhaust area (AJ) with fan rotor corrected speed -^2 at corresponding constant rotor speed values. With existing control systems, a deceleration from intermediate thrust to low thrust (B curve 146 ) with little or no increase in exhaust area(AJ) results in a large reduction in low rotor (N-|) speed. Throttle response is adversely effected since the -] rotor speed must substantially increase in order for engine thrust to subsequently return to intermediate thrust (A). The large rotor speed excursion reduces the rotor low cycle fatigue (LCF) life drastically. Also, the smaller low power exhaust area reduces low speed compres- sor stability.

In contrast, an active geometry control system of the present invention schedules the exhaust area to open syn- . chronously with other engine parameters. As indicated by curve 148, identical levels of low thrust can be obtained at higher low rotor (N-j) speeds (C) . With a control system of the present invention rotor speeds at low thrust are higher than are obtained with existing control systems since thrust level is not only dependent upon a reduction in rotor speed but rather is significantly effected by the larger exhaust area. Therefore, an active geometry control system of the present invention provides dramatic improvement in thrust response because the engine spool speeds do not have to change significantly. The reduced rotor speed excursions correspondingly increase rotor low cycle fati-0 gue(LCF) life. Moreover, the larger exhaust area at low thrust levels enhances compression systems stability. The larger exhaust area (AJ) reduces back pressure on the low pressure compressor by means of a bypass duct. Since the low pressure turbine is "un-choked" at low power, the correspon-ding reduction in nozzle pressure lowers the high pressure compressor operating line and hence enhances high compressor stability.

The active geometry control system provided by the present invention utilizes the concept of sychronous schedu-ling of all gas path variable parameters and gas generator (main engine combustor) fuel flow parameter. The synchronization of the fuel flow rate limit during acceleration and the HCVV(req) rate limit provides unique control and adjustment capability of high compressor operating line. If the engine reacceleration command is performed before the expiration of the period set by the deceleration timer and the control system has not permitted scheduling of steady state high compressor variable vane position extremely fast thrust accelerations are possible since power is not dependent up-on significant rotor speed acceleration.

As mentioned above this invention is particularly efficacious for military aircraft powered by a twin spool, axial flow fan jet. A typical installation where this invention would.be applicable is for the F-100 family of engines manu- factured by Pratt & Whitney Aircraft, a division of United Technologies Corporation, the assignee of this patent application, and reference should be made thereto for further details.

To fully appreciate the objectives of this invention reference should be made to FIG 8 which shows a typical operating line of the LCVV plotted against N-| (unless otherwise specified all speed parameters are intended to be their corrected value) . Since the fan and low pressure compressor are a unitary rotor and connected by the same shaft, the fan and low compressor speed are obviously identical. Curve A illustrates the usual operating line of the low speed compressor and, as noted, Fn or Wa increases as increases. Wa is corrected airflow parameter normalized as a function of total inlet temperature and pressure. In this particular , the high power conditions, close at low power condition and are modulated therebetween.

The operating line (Curve A) is manifested by the ccor-dinated efforts of controlling fuel and the vanes which are typically controlled as a function of engine and aircraft operating variables. According to this invention rapid transients can be produced by controlling the LCVV such that low compressor airflow is modulated at constant or nearly con-stant -| . Curve 3 is illustrative of a transient deceleration to a lower power thrust value identified as point C on Curve B. As is apparent from the graph, the active low compressor control (ALCC) serves to hold N-j constant (Curve B) while modulating the LCW until the engine attains the tar-get (point C) . If a bodie is exercised, the engine is accelerated back to the high power condition (point E) by again merely adjusting LCW and controlling along curve B.

If the bodie maneuver in the above example is not exercised, would remain fixed at point C. In accordance with this invention, after a given interim the ALCC trims the LCW and fuel flow to return the setting to the operating line (Curve A) to the steady state operating point illustrated by point G along the constant Wa line (Curve H in this example) . This is the condition of optimum engine perfor-mance from a TSFC and stability standpoint.

While the example above describes a transient excursion at a selected speed setting, these excursions will occur at any point along the operating line.

In its preferred embodiment the fuel control and ALCC are the electronic, digital type of controller and the fuel control may be, for example, the fuel control Model # EEC- 104 manufactured by the Hamilton Standard division of United Technologies Corporation (incorporated herein by reference) or may be implemented by other media, such as, hydro-me- chanical, electro-mechanical, and the like. As will be best understood by those skilled in the art, once the function's logic is understood, state-of-the-art technology can be implemented to execute this invention.

As is apparent from the foregoing, the control will operate to attain tha gas generator (N2 ) operating line by normal adjustment by the fuel control by the flew of fuel to the engine's combustor. Normal transients are likewise manifested in this manner. For rapid transient excursions, such as those contemplated when operating in the combat box, the control will automatically control the engine's acceleration and deceleration modes as a function of the primary control parameter N-j . thetefore, is scheduled as a function of a power lever angle a which automatically sets engine thrust by proper scheduling of total engine airflow and engine pres-sure ratio by virtue of adjusting the LCW. Thus, in this instance, N-j is utilized as the primary control parameter for scheduling both fuel flow and LCW.

The type of engine for which this invention is particularly efficacious is schematically illustrated in FIG 9 as being a fan jet engine generally illustrated by reference numeral 110 and comprised of an axial flow twin spool configuration. As schematically shown, the high pressure spool consists of a plurality of stages of compressors generally illustrated by reference numeral 112 driven by the first stage turbine 114 and interconnected thereto by shaft 116.

The low pressure spool consists of the fan/low pressure compressor combination generally illustrated by reference numeral 118 where the fan portion discharges through the outer annular passageway 120 and the low pressure compres-sor discharges into the inlet of the high pressure compressor 112.·. Low pressure turbine stages 122 serve to. power the fan/ low pressure compressor 118 which is connected thereto by shaft 124. A suitable combustor 123 is interposed between the compressor section and turbine section where fuel is combusted to provide the working medium for powering the turbines and generating thrust. Fuel is fed to the combustor by throttle valve 125 as will be described hereinbelow.' The engine may utilize a suitable augmentor generally indicated by reference numeral 126 and a suitable variable jet nozzle 128. The engine has variable vanes 130 at the inlet of the fan and may also include variable vanes 136 in the high pressure compressor section.

As an understanding of the details of the engine are not necessary for an understanding of this invention, for the sake of cenvenience and simplicity, they are omitted herefrom. It is, however, necessary to understand that the invention is applicable in a single or multiple spool engine that has variable vanes at the inlet of the fan and/or low pressure compressor. . ·-■ The electronic digital controller is comprised of four distinct circuits', W, X, Y and Z. While each of these circuits respond to a plurality of measured variables, these signals are interconnected to each of the circuits so that these signals are shared as needed.

Circuit W serves to derive a target thrust signal in accordance with the input of the power lever by scheduling engine airflow as a function of power lever position (o , r an inlet pre,,ssure^and tejnperature and aircraft Mach No. The output of the function generator 140 become Tfie" "input to function generator 142 for setting a N-| request signal. This -| request signal as applied to function generator 144 which becomes the control parameter for scheduling a . reference LCW position (LCW ref ) . The LCVV ref signal and the LCVV trim signals are applied to summer 145 and this value is the target value of the LCVV. The LCW target value is restricted by a minimum and maximum limiter by the function generators 146 and 148 respectively as a function of actual corrected to inlet conditions. The output of summer 1 5 which the LCVV target sets is the LCW request signal (LCW req) . As shown in block 149 in the block diagram, the LCVV req signal may be rate limited by function generator 151 as a function of altitude, Mach No., PT2 and/ or TT2 · It is apparent from the foregoing that in the example in FIG 8, the target signal C which is the output of the summer 145 is attained by adjusting LCW to attain the proper Wa value while maintaining -j constant (scheduling along Curve B) .

In circuit X, the accel/decel selector logic 150 locks a value of Ni as a function of the motion of the power le- er and is compared to the measured value of N^ by summer 152 to provide an N-| error signal. This difference is the input for a proportional-plus-integral controller 154 which is applied to summer 156 to trim LCVV to adjust N-| to the N1 lock value (Curve B of FIG 8) .

The decel timer 156 in circuit Y is activated by the decel signal from the accel/decel selector 150. Values of the timer 156 are set as a function of altitude and Mach No. At the termination of the signal produced by timer 156, the LCVV's are slewed from their temporary position to the steady state schedule (LCWSS) , scheduled by function generator 160 as a function of actual N- via the elimination of LCVVERR, a function of timer 156 setting the LCW(req) and LCWSS. The summer 166 schedules the LCW steady state signal as a function of the LCVV vane position via feedback line 168. This function reduces N-j to the normal steady state operating speed (point G in the example in FIG 8) .

Fuel flow is regulated by circuit Z which schedules fuel ■flow by generating a W /PB signal as a function of the request signal generated by the function generator 143 in the W circuit (where is fuel flow in pounds per hour and Pg is the high compressor discharge pressure or the pressure of the engine's combustor).

The function generator 170 serves to limit the Wj/PB maximum and minimum values as a function of measured 2 cor- rected to the inlet value. The Wf/PB may be rate limited as shown by the function generator 172 as a function of any number of engine and/or aircraft operating variables , such as, Alt., Mach No.,PT2 and TT2. The Wf/PB request value is then multiplied by a suitable multiplier 174 by the measured Ρβ value to produce a fuel flow (Wf) signal for driving the throttle valve 125 and regulating fuel flow to, the engine's combustor .

The synchronization of the fuel flow rate of change and the LCW(req) rate of change provide unique control and ad- justment capability of the low pressure compressor operating line, especially during engine thrust decels. If the engine has been previously at high power and a thrust re-accel performed before the decel timer permits scheduling of LCWSS, extremely fast thrust accelerations are possible since power is not dependent on low rotor speed re-acceleration.

As mentioned above this invention is particularly efficacious for military aircraft powered by a twin spool, axial flow gas turbine engine either of the fan or straight jet variety. A typical installation where this invention v.-ould be applicable is for the F-100 family of engines manufac- of the engine's combustor. jrhe func^oj^jqenerator 170 serves to limit the Wf/PB maximum and minimum values as^ function _of_ measured N2 cor- rected to the inlet value. The Wf/PB may be rate limited as shown by the function generator 172 as a function of any number of engine and/or aircraft operating variables, such as, Alt., ach No.,PT2 and TT2. The Wf/PB request value is then multiplied by a suitable multiplier 174 by the measured Ρβ value to produce a fuel flow (Wf) signal for driving the throttle valve 125 and regulating fuel flow to the engine's formed before the decel timer permits scheduling of LCVVSS , extremely fast thrust accelerations are possible since power is not dependent on low rotor speed re-acceleration.

As mentioned above this invention is particularly efficacious for military aircraft powered by a twin spool, axial flow gas turbine engine either of the fan or straight jet variety- A typical installation where this invention would e applicable is for the F-100 family of engines manufac- tured by Pratt & Whitney Aircraft, a division of United Technologies Corporation, the assignee of this patent application, and reference should be made thereto for further details .

To fully appreciate the objectives of this invention reference should be made to FIG 10 which shows a typical operating line of the 'gas generator, i.e. , HCW plotted against N2 (unless otherwise specified all speed parameters are intended to be their corrected value) . Curve A illus-trates the usual operating line of the gas generator and, as noted, power or Fn increases as N2 increases. In this particular plot showing the relationship of the HCVV, the vanes open at the high power conditions, close at low power condition and are modulated therebetween.

The operating line (Curve A) is manifested by the coordinated efforts of controlling fuel and the vanes which is typically controlled as a function of engine and aircraft operating variables. According to this invention rapid transients can be produced by controlling the HCW such that the high compressor airflow is modulated at constant or nearly-constant Curve B is illustrative of a transient deceleration to a lower power thrust value identified as point C on curve B. As is apparent from the graph, the active high comp'ressor control (AHCC ) serves to hold N2 constant (curve B) while modulating the HCVV until the engine attains the target (point C) . If a bodie is exercised, the engine is accelerated back to the high power condition v(point E) by again merely adjusting HCW and controlling along curve B.

If the bodie maneuver in the above example is not exer- cised, N2 would remain fixed at point C. In accordance with this invention, after a given interim the AHCC trims the HCW and fuel flow to return the setting to the operating line (curve A) to the steady state operating point illustrated by point G along the line, N-]=4500RPM in this example. This is the condition of optimum engine performance from a TSFC and stability standpoint.

While the example above describes a transient excursion tured by Pratt & Whitney Aircraft, a division of United Technologies. Corporation , the assignee of this patent application, and reference should be made thereto for further details .

To fully appreciate the objectives of this invention reference should be made to FIG 8 which shows a typical operating line of the gas generator, i.e., HCW plotted against 2 (unless otherwise specified all speed parameters are intended to be their corrected value) . Curve A illus-trates the usual operating line of the gas generator and, as noted, power or Fn increases as 2 increases. In this particular plot showing the relationship of the HCVV, the vanes open at the high power conditions, close at low power condition and are modulated therebetween.

The operating line (Curve A) is manifested by the coordinated efforts of controlling fuel and the vanes which is typically controlled as a function of engine and aircraft operating variables. According to this invention rapid transients can be produced by controlling the HCW such that the high compressor airflow is modulated at constant or nearly constant 2 · Curve B is illustrative of a transient deceleration to a lower power thrust value identified as point C on curve B. As is apparent from the graph, the active high compressor control (AHCC ) serves to hold N2 constant (curve B) while modulating the HCVV until the engine attains the target (point C) . If a bodie is exercised, the engine is accelerated back to the high power condition (point E) by again merely adjusting HCVV and controlling along curve B.

If the bodie maneuver in the above example is not exer-cised, N2 would remain fixed at point C. In accordance with this invention, after a given interim the AHCC trims the HCVV and fuel flow to return the setting to the operating line (curve A) to the steady state operating point illustrated by point G along the N1 line, N1=4500RPM in this example. This is the condition of optimum engine performance from a TSFC and stability standpoint.

While the example above describes a transient excursion at a selected N2 speed setting, these excursions will occur at any point along the operating line.

In its preferred embodiment the fuel control and AHCC are the electronic, digital type of controller and the fuel control may be, for example, the fuel control Model EEC- 106 manufactured by the Hamilton Standard division of United Technologies Corpora't ion ( incorporated herein by reference) or may be implemented by other media, such as, hydro-mechanical, and the like. As will be best understood by those skilled in the art, once the function's logic is understood, state-of-the-art technology can be implemented to execute this invention .

As is apparent from the foregoing, the control will operate to attain the gas generator (N2) operating line by normal adjustment by the fuel control by the flow of fuel to the engine's combustor. Normal transients are likewise manifested in this manner. For rapid transient excursions, such as those contemplated when operating in the combat box, the control will automatically control the engine's acceleration and deceleration modes as a function of the primary control parameter N-| . N-j, therefore, is scheduled as a function of power lever angle <=_. which automatically sets engine thrust by proper scheduling of total engine airflow and engine pressure ratio by virtue of adjusting the HCW and fuel flow. Thus, in this instance, -j is utilized as the primary control parameter for scheduling both fuel flow and HCW.

The type of engine for which this invention is particularly efficacious is schematically illustrated in FIG 11 as being a fan jet engine generally illustrated by reference numeral 210 and comprised of an axial flow twin spool configuration. As schematically shown, the high pressure spool consists of a plurality of stages of compression generally illustrated by reference numeral 212 driven by the first stage turbine 214 and interconnected thereto by shaft 216.

The low pressure spool consists of the fan/low pressure compressor combination generally illustrated by reference numeral 218 where the fan portion discharges through the at a selected 2 speed setting, these excursions will occur at any point along the operating line.

In its preferred embodiment the fuel control, and AHCC are the electronic, digital type of controller and the fuel control may be, for example, the fuel control Model^ EEC-106 manufactured by the Hamilton Standard division of United Technologies Corporation ( incorporated herein by reference) or may be implemented by other mediums, such as, hydro-mechanical, and the like. As will be best understood by those skilled in the art, once the function's logic is understood, state-of-the-art technology can be implemented to execute this invention.

As is apparent from the foregoing, the control will operate to attain the gas generator (N2) operating line by normal adjustment by the fuel control by the- flow of fuel to the engine's combustor. Normal transients are likewise manifested in this manner. For rapid transient excursions, such as those contemplated when operating in the combat box, the control will automatically control the engine's acceleration and deceleration modes as a function of the primary control parameter 1. , therefore, is scheduled as a function of power lever angle di which automatically sets engine thrust by proper scheduling of total engine airflow and engine pressure ratio by virtue of adjusting the HCW and fuel flow. Thus, in this instance, is utilized as the primary control parameter for scheduling both fuel flow and HCW.

The type of engine for which this invention is particularly efficacious is schematically illustrated in FIG 11 as being a fan jet engine generally illustrated by reference numeral 210 and comprised of an axial flow twin spool configuration. As schematically shown, the high pressure spool consists of a plurality of stages of compression generally illustrated by reference numeral 212 driven by the first stage turbine 214 and interconnected thereto by shaft 216.

The low pressure spool consists of the fan/low pressure compressor combination generally illustrated by reference numeral 218 where the fan portion discharges through the outer annular passageway 220 and the low pressure compressor discharges into the inlet of the high pressure compressor. Low pressure turbine stages 222 serve to power the fan low pressure compressor 218 which is connected thereto by shaft 224. A suitable combustor 223 is interposed between the ■ compressor section and turbine section where fuel is combusted to provide the working medium for powering the turbines and generating thrust. Fuel is fed to the combustor by throttle valve 225 as will be described hereinbelow. The engine may utilize a suitable augmentor generally .indicated by reference numeral 226 and a suitable variable jet nozzle 228. The engine may also be equipped with variable vanes 230 and 236 at inlets of the fan and the high pressure compressor respectively.

As an understanding of the details of the engine are not necessary for an understanding of this invention, for the sake of convenience and simplicity, they are omitted herefrom. It is, however, necessary to understand that the invention is solely applicable in a single or multiple twin spool engine that has variable vanes at the inlet and/or other stages of the high pressure compressor.

The electronic digital controller is comprised of four distinct circuits, W, X, Y and Z. While each of these circuits respond to a plurality of measured variables, these signals are interconnected to each of the circuits so that these signals are shared as needed.

In circuit W engine airflow is scheduled as a function of power lever position oL, fan inlet pressure and temperature and aircraft Mach No. The output of the function gene- rator 240 becomes the input to function generator 242 for setting a request signal. This N-j request signal becomes the control parameter for scheduling both a reference HCVV position (HCWREF) and a HCW error signal (DHCW) . The DHCW signal is derived by applying a gain (slope of the N-j line of FIG 10) from gain controller 244 and a 2 speed error signal (DEBASE) which is generated as an error signal between a referenced N2 preselected value (N-,Ref) and a locked N2 value (curve B in. the example of FIG 10)set by the accel/ outer annular passageway 220 and the low pressure compressor discharges into the inlet of the high pressure compressor. Low pressure turbine stages 222 serve to power the fan/ low pressure compressor 218 which is connected thereto by shaft 224. A suitable combustor 223 is interposed between the compressor section and turbine section where fuel is combusted to provide the working medium for powering the turbines and generating thrust. Fuel is fed to the combustor by throttle valve 225 as will be described hereinbelow. The engine may utilize a suitable augmentor generally indicated by reference numeral 226 and a suitable variable jet nozzle 228. The engine may also be equipped with variable vanes 230 and 236 at inlets of the fan and the high pressure compressor respectively.

As an understanding of the details of the engine are not necessary for an understanding of this invention, for the sake of convenience and simplicity, they are omitted . herefrom. It is, however, necessary to understand that the invention is solely applicable in a single or multiple twin spool engine that has variable vanes at the inlet and/or other stages of the high pressure compressor.

The electronic digital controller is comprised of four distinct circuits, W, X, Y and Z. While each of these circuits respond to a plurality of measured variables, these signals are interconnected to each of the circuits so that these signals are shared as needed.

In circuit W engine airflow is scheduled as a function of power lever position o , fan inlet pressure and temperature and aircraft Mach No. The output of the function gene- rator 240 becomes the input to function generator 242 for setting a request signal. This ·| request signal becomes the control parameter for scheduling both a reference HCVV position (HCWREF) and a HCW error signal (DHCW) . The DHCW signal is derived by applying a gain (slope of the line of FIG 10) via gain controller 44 and a speed error signal (DEBASE) which is generated as an error signal between a referenced 2 preselected value (^Ref) and a locked 2 value (curve B in the example of FIG 1,0)aet by the accel/ decel selector 248 in circuit X of the AHCC . The .2 lock assures that the power lever position will consistently attain a given thrust for each power lever angle and won't be adversely affected by power extraction or changes in the engine's performance due to wear and tear. As noted, the N2 lock is a mixture of the power lever angle signal and actual 2- The targeted value, i.e., the desired thrust corresponding to the position of the power lever, is a sum of DHCVV, HCWREF and HCVV trim. The HCVV trim signal is a speed error signal (D 2) from the summer 250 that adds the N2 locked value and the actual 2. The DN2. signal is fed to a proportional and integral controller 252 that trims the 2 speed signal to assure it maintains a constant or substantially constant value when the engine decelerates to the targeted value (curve B of example in FIG 10) .

A maximum and minimum HCVv) value set by function generators 254 and 256 limits the HCVV base signal as a function of measured 2- The limits regulate the opening and closing of vanes of the HCVV to stay within prescribed boundaries of the N2 compressor map. If necessary and to assure that the vanes do not open or close too quickly, a rate limiter 257 which may be adjusted as a function of any number of engine and/or aircraft operating parameters such as altitude, Mach No., PT2 and TT2 applied to the output of summer 259.

Circuit Y of the controller serves to return the engine to its operating line (curve A of FIG 10) after a predetermined time after the target value has been reached. The time set in a timer 260 may be preselected as a function of altitude and/or aircraft Mach No . At the termination of the timer, the HCVV are slewed from thei temporary positions to the steady state schedule (HCVVSS) which is established by the function generator 262 as a function of measured N2. The controller 264 is a function of HCVVSS, a feedback signal of the position of HCVV and the slew re-quest signal (the output of timer 260) . As HCVV and Wf are trimmed, the N2 value is returned to the steady state operating line (point G in the example shown in FIG 10) .

- - Fuel flow is regulated by circuit Z which schedules fuel flow by generating a Wc/Pg signal as a function of the request signal generated by the function generator 242 in the circuit (where is fuel flow in pounds per hour and Pg is the high compressor discharge pressure or the pressure of the engine, ' s combustor ) .

The function generator 270 serves to schedule the Wf/ Pg value as a function of measured The W^/Pg may be rate limited as shown by the function generator 272 as a function of any number of engine and/or aircraft operating variables, such as, Alt., Mach No., PT7 and TT2 · The W^/Pg request value is then multiplied by a suitable multiplier 274 by the measured Pn value to produce a fuel flow (Wc) signal for driving the throttle valve 225 and regulating fuel flow to the engine's combustor.

The synchronization of the fuel flow rate of change and the HCVV(req) rate of the change provide unique control and adjustment capability of the high compressor operating line, especially during engine thrust accels. If the engine has been previously at high power and a thrust re-accel performed before the decel timer permits scheduling of HCWSS, extremely fast thrust accelerations are possible since power is not dependent on significant high rotor speed re-acceleration, if any.

Claims (19)

91682/3 C L A I M S

1. For a gas turbine engine for powering aircraft having a pair of spools comprising a high pressure compressor/turbine combination and a fan/low pressure compressor and turbine combination solely aerodynamically coupled, a burner for combusting fuel and air for powering said turbines, a power lever, fuel regulating means for regulating the flow of fuel to said burner and variable area vanes for regulating the air flow to said fan/low pressure compressor, means for controlling said engine to operate on a low pressure compressor steady state operating line, the improvement comprising control means responsive to the position of said power lever to regulate rapid accelerations and decelerations in response to a preselected low pressure compressor airflow parameter for controlling said variable vanes and said fuel regulating means, first means responsive to a plurality of engine and aircraft variables for generating a signal indicative of a desired low pressure compressor speed, means responsive to the difference between power lever position and the actual low pressure compressor speed including a proportional and integral controller for adjusting said variable vanes to accelerate or decelerate said engine along a constant low pressure compressor speed to the targeted value selected by said power lever position, and additional control means responsive to said airflow parameter for controlling said fuel regulating means and said variable vanes to return said engine to operate on said low pressure compressor steady state operating curve and timer means for actuating said additional control means upon reaching a predetermined time interval. 31 91682/3

2. For a gas turbine as claimed in claim 1 wherein said low pressure compressor speed parameter is a function of power lever position and an engine operating variable.

3. For a gas turbine engine as claimed in claim 2 wherein said engine operating variable is total pressure measured at the inlet of said fan/low pressure compressor.

4. For a gas turbine engine as claimed in claim 2 wherein said engine operating variable is total temperature measured at the inlet of said fan/low pressure compressor.

5. For a gas turbine engine as claimed in claim 2 wherein said low pressure speed parameter is also a function of an aircraft operating variable.

6. For a gas turbine engine as claimed in claim 5 wherein said aircraft operating variable is Mach No.

7. For a gas turbine engine as claimed in claim 6 wherein said aircraft operating variable is altitude.

8. For a gas turbine engine as- claimed in claim 1 wherein said low pressure compressor speed parameter is a function of aircraft Mach No. and altitude, engine total pressure and temperature measured at the inlet of said fan/low pressure compressor and pilot lever position.

9. For a gas turbine engine as claimed in claim 8 wherein said time interval is made to vary as a function of aircraft altitude and Mach No. 32 91682/3

10. For a gas turbine engine as claimed in claim 9 including means responsive to measured low pressure compressor speed to limit the minimum and maximum position of said variable vanes.

11. For a gas turbine engine as in claim 10, wherein said, control for said fuel regulating means is responsive to a function of request and measured high pressure compressor speed for generating a W^/Pg signal and means for multiplying said Wf/PB signal and measured burner pressure for generating a signal for controlling said fuel regulating means.

12. A control for controlling a gas turbine engine for powering aircraft during rapid acceleration and rapid deceleration operating modes, said gas turbine engine having at least one compressor and turbine driving said compressor, a burner, fuel flow means regulating the flow of the fuel to said burner, a power lever, control means responsive to said power lever for controlling said engine in accordance with a predetermined operating line of said compressor, variable geometry control means to control the airflow to said compressor, means for generating a target value that is outside of said predetermined operating line of said compressor, means for rapid acceleration and deceleration operating modes responsive to power lever position to generate a first signal, means responsive to said first signal for setting a fixed speed of said compressor for accelerating or decelerating said compressor to said target value, means responsive to an engine parameter to provide a second signal after a predetermined time elapsed to simultaneously adjust the fuel flow means and said variable geometry control means to trim the speed of said compressor from said targeted value back to said predetermined operating line of said compressor. 33 91682/3

13. A control for controlling a gas turbine engine for powering aircraft during rapid acceleration and deceleration operating modes, said gas turbine engine having a first spool comprising a high pressure compressor and a first turbine driving said high pressure compressor and a second spool comprising a low pressure compressor and a second turbine driving said low pressure compressor and said first spool and said second spool being solely aerodynamically coupled, fuel metering means for regulating the flow of fuel to said engine, and variable vanes for said high pressure compressor for regulating the airflow thereto, means for controlling said engine to operate on a predetermined high compressor steady state operating curve and having the speed of said first spool and said second spool being at a predetermined relationship, a power lever, said control including means responsive to the position of said power lever for establishing a thrust value for adjusting said engine to accelerate and decelerate rapidly, said control having additional control means responsive to engine and aircraft operating variables for generating a first signal requesting a desired corrected low compressor speed (N^), means responsive to said first signal and a preselected fixed corrected high pressure compressor speed (N2) for generating a desired airflow from said variable vanes, means responsive to said power lever for selecting a constant N2 , means responsive to the difference of said selected constant N2 and measured N2 including a proportional and integral controller for adjusting said variable vanes to accelerate or decelerate said engine to said value established by said power lever, and means responsive to a timed signal to adjust both said fuel metering means and said variable vanes as a function of said desired corrected low pressure speed to return said engine to operate on said high compressor steady state operating curve. 91682/ 3

14. 1 . A control as in claim 13 including means responsive to measured N2 for establishing a maximum and minimum value of air flowing through said variable vanes by limiting the opened and closed position of said variable vanes .

15. A control as in claim 14 wherein said engine and aircraft operating variables for said additional control means includes aircraft Mach No. and total pressure at the inlet of said high pressure compressor.

16. A control as in claim 15 including the additional engine operating parameter of total temperature at the inlet of said high pressure compressor.

17. A control as in claim 16 including means responsive to aircraft altitude and Mach No. and total pressure and total temperature measured at the inlet of said high pressure compressor for limiting the rate of travel of said variable vanes.

18. A control as in claim 17 wherein said engine includes a burner and wherein said means responsive to said timed signal for adjusting fuel flow includes means responsive to a function of v said requested and measured N2 for producing a second signal (Wj/Pg) and means responsive to the multiplicand of Wf/Pg an<^ measured pressure of said burner for controlling said fuel metering means .

19. A control as in claim 14 wherein said time signal is modified as a function of aircraft altitude and Mach No. For the Applicant, Co. 35