FUEL CONTROL WITH SMOOTH MODE TRANSITION

Abstract: A gas turbine engine control system which regulates fuel flow based upon a schedule which is a function of a primary parameter, such as engine pressure ratio, switches to a schedule based upon a secondary parameter, such as compressor speed, upon the inability to function in the primary parameter mode. Upon switchover to the secondary parameter mode the scheduled value of the secondary parameter is down-trimmed to make it equivalent to the last known good value of the primary parameter so that the engine will not experience a sudden change in power level upon switchover. The down-trim is limited. When the primary mode failure occurs during a transient the down-trim is limited to the difference between a scheduled reference value of the secondary parameter and a value of the secondary parameter calculated immediately preceeding loss of the primary parameter mode and based upon a known relationship between the primary and secondary parameter. The down-trim limit assures a smooth transition from the primary to the secondary control mode when switchover occurs during an engine transient.

Description: Technical Field This invention relates to fuel controls for gas turbine engines. Background Art Fuel controls for gas turbine engines operate in a closed loop fashion based on an engine parameter, such as engine pressure ratio (EPR) which is the ratio of the engine exhaust to engine inlet total pressure P₅/P₂. A desired or reference EPR is computed based upon throttle setting and prevailing atmospheric conditions and is compared to the actual EPR of the engine; fuel is modulated to drive the resultant error term to zero. A detected failure of any one of the signals necessary to compute either the reference EPR or the actual EPR forces operation of the control based upon an alternate control mode, such as the speed N₁ of the low compressor if the engine is a twin spool gas turbine engine. In that case, a reference N₁ (N 1ref ) is computed based upon prevailing conditions and throttle setting. This is compared to the actual N₁ and fuel is modulated to drive the resultant error term to zero. Because the characteristics of these different references are different, there could be a sudden change in engine speed at the time of transfer. This is sometimes referred to as a "bump". Bumps, at a minimum, can be unsettling to the pilot and passengers of the aircraft. Furthermore, if they occur during transient engine operation they can be dangerous, such as when the pilot suddenly calls for full power, and the engine does not respond in the manner expected. In the PW2037 twin spool engine manufactured by the Pratt & Whitney Division of United Technologies Corporation, it is known to eliminate bumps when a control mode change occurs by attempting to force the new reference parameter to match the failed reference parameter at the instant of failure. More specifically, if EPR is the primary mode and N₁ is the backup mode, it is known to continuously monitor N₁ during EPR operation and to use the last value thereof prior to EPR failure to trim the reference speed schedule such that N 1ref (trimmed) equals the last measured value of N₁ at the instant of EPR failure. If the desired N₁ just after EPR failure (i.e. trimmed N 1ref ) is almost the same as N₁ just before the failure, the bump is virtually eliminated. This will be the case when EPR fails during steady state engine operation or at low rates of acceleration or deceleration; however, if EPR failure occurs near the onset of a significant engine transient the bump will not be eliminated as a result of the great difference between the engine speed at the time of EPR failure and the desired engine speed (N 1ref ) based upon EPR at the new throttle setting at the time of EPR failure. The fault logic will therefore overcompensate and the result may be a severe loss of engine thrust. Disclosure of the Invention One object of the present invention is a fuel control for a gas turbine engine which can transfer from operation based upon a first engine parameter to operation based on a different engine parameter with minimum effect on engine operation at the time of switchover. Another object of the present invention is a fuel control for a gas turbine engine which transfers from one mode of operation to another mode of operation with minimal effect on engine operation at the time of switchover, even if switchover occurs in the midst of an engine acceleration or deceleration. According to the present invention, in a gas turbine engine control system which regulates fuel flow on the basis of a primary parameter schedule and switches to a secondary parameter schedule upon the inability to reliably determine the primary parameter, the control continuously calculates the value of the secondary parameter at which it would be desired to operate the engine based upon a known relationship between the first and second parameter, the difference between a scheduled reference value of the second parameter and such calculated value of the second parameter immediately preceeding loss of the first parameter being used to limit the down-trim of the scheduled reference value of the second parameter upon loss of the first parameter. With this control technique the down-trim limit is always calculated using an estimate of the desired value of the second parameter at the new throttle setting based upon the primary parameter schedule. If primary parameter failure occurs during a transient, this estimated value may be considerably different from the actual value of the secondary parameter at the instant of primary parameter failure, (particularly if the secondary parameter changes relatively slowly during transients) and will be much closer to the value of the secondary parameter which the new throttle setting would have called for had the primary parameter not been lost. In a preferred embodiment the primary parameter is EPR and the secondary parameter is engine speed. When EPR failure occurs the control switches to an engine speed schedule to control the fuel flow. The speed schedule is not identical to the EPR schedule and, if not corrected, would result in a sudden engine speed change upon switchover, even without throttle movement. In the control of the present invention the speed schedule is trimmed upon EPR failure by an amount equal to the difference between the scheduled and actual value of engine speed just prior to EPR failure to compensate for the difference between the EPR and speed schedules at the current throttle setting. The same engine speed will therefore be called for by the control at the instant the operating mode is switched. This trim amount is limited by a trim value which is calculated using an estimate of the desired engine speed at the current throttle setting based upon the EPR schedule in use prior to EPR failure, rather than being calculated using the actual engine speed at the time of EPR failure. This is done because the actual engine speed in the midst of a transient might be much lower than the speed which the EPR schedule would have called for at the new throttle setting. Without this limit the result could be excessive trim when EPR failure occurs during a transient. The foregoing and other objects, features and advantages of the present invention will become more apparent in the light of the following detailed description of preferred embodiments thereof as illustrated in the accompanying drawing. Brief Description of the Drawing Fig. 1 is a schematic and block diagram of a twin spool gas turbine engine incorporating the control system of the present invention. Fig. 2 is a schematic and block diagram showing another aspect of the control system of the present invention. Best Mode for Carrying Out the Invention As an exemplary embodiment of the present invention, consider the twin spool turbofan gas turbine engine shown in the drawing and generally represented by the reference numeral 10. The engine comprises a low compressor 12 connected through a shaft to a low turbine 14; a high compressor 16 connected through a shaft to a high turbine 18; and a burner section 20 disposed between the high compressor and the high turbine. A plurality of fuel nozzles 22 spray fuel into the burners 24 of the burner section 20. Fuel flow rate into the nozzles 22 is varied by a valve 26. An electronic engine control automatically regulates engine operation, such as fuel flow rate, based upon pilot demand (throttle setting), various aircraft and engine parameters and scientific and empirically developed relationships between various parameters. In this embodiment the electronic engine control's primary mode of fuel flow control is based upon engine total pressure ratio (EPR), which is the ratio of the engine outlet pressure (P₅) to the engine inlet pressure (P₂). As shown in the drawing, a signal 28 indicative of the inlet pressure and a signal 30 indicative of the outlet pressure are delivered to a divider 32. The output signal 34 from the divider is the engine pressure ratio, which signal is delivered to a portion of the fuel control designated by the box 36. Within the control portion 36 is a schedule of engine pressure ratios against which the actual pressure ratio is compared. The control portion 36 sends a signal 38 to the fuel nozzle valve 26 to adjust the fuel flow, and thereby the engine speed, until the actual engine pressure ratio matches the scheduled engine pressure ratio for the particular throttle setting. If either the primary fuel control parameter, engine pressure ratio, or the engine pressure ratio schedules become unavailable or unreliable (hereinafter either occurrence is referred to as an EPR mode failure), the control portion 36 switches to a secondary parameter, low compressor speed N₁, to control the fuel flow. With reference to the drawing, a control portion 40 continuously produces an output signal 42 indicative of a scheduled low compressor reference speed N 1ref . This reference speed is the maximum speed that the engine can tolerate under existing conditions and throttle setting. A signal 44 indicative of the actual low compressor speed N₁ is subtracted from N 1ref in a subtractor 46; and a signal 48 indicative of that difference is delivered to a switch 50. Of course, while the engine is still being controlled based upon EPR, N₁ will be a value which is a function of EPR in accordance with the EPR schedule built into the control. A signal 52 is delivered to the switch 50 when an EPR mode failure occurs. When no signal 52 is present the output 53 of the switch 50 is the presently calculated value of the signal 48, which is sent to a select-low gate 54. When the switch 50 receives the signal 52, its output 53 is a signal 56 which is indicative of the value of the signal 48 last calculated prior to the switch receiving the signal 52; and that value of the signal 48 is thereafter continuously delivered to the select low gate 54. A function generator 58 continuously receives a signal 60 indicative of the aircraft Mach number M n and a signal 61 from the control portion 36 indicative of the scheduled engine pressure ratio. Based upon these inputs the function generator 58 generates a signal 62 (N E C₂) indicative of the low compressor speed corrected to the temperature at the low compressor inlet. In a manner well known in the art, the corrected low compressor speed 62 and an appropriate multiplier 64 (a function of temperature at the low compressor inlet) are delivered into a multiplier 66 to yield a signal 68 indicative of an estimated value of the low rotor speed (N E ). The estimated low rotor speed signal 68 is subtracted from N 1ref in a subtractor 70 and a signal 72 indicative of that difference is delivered to a switch 74. A signal 76 is delivered to a switch 74 when an EPR mode failure occurs. When there is no signal 76 the switch 74 passes the presently calculated value of the signal 72. Upon receiving the signal 76 the output 78 of the switch 74 has the value of a signal 80 which is indicative of the value of the signal 72 last calculated prior to the switch 74 receiving the signal 76. The value of the signal 78 therefore represents the difference between N 1ref and an estimated low rotor speed based upon an engine pressure ratio schedule and the current engine throttle setting, and not based upon the current low rotor speed N₁. Thus, if the engine pressure ratio or engine pressure ratio schedules become unavailable or unreliable immediately after the throttle has been moved and before the engine has had an opportunity to change from its old speed to its new speed, the signal 68, which is the estimated compressor speed N E , will be approximately the speed normally called for by the engine pressure ratio schedule at the new throttle setting. In contrast, the speed signal 44 will be the actual engine speed at the time of EPR mode failure, which may be considerably different from the desired speed based upon the throttle setting and the EPR schedule. The speed difference signals 78 and 53 are both fed to the select-low gate 54, and the lower of the two signals, represented by the output signal 82, is passed to a select-high gate 84. The high gate 84 receives the signal 82 as well as a signal 86. The signal 86 is always zero. Thus, if the signal 82 is negative, the output signal 88 from the high gate 84 will be zero; otherwise it will be identical to the signal 82. The high gate 84 is used to assure that the value of the signal 88 (the N 1ref trim) is never negative, which could happen when the EPR mode failure occurs during a deceleration. A negative value for the signal 88 would result in a value of the signal 92 which is higher than N 1ref , which cannot be allowed since N 1ref is the maximum low compressor speed which the engine can safely tolerate. Thus, only a down-trim is applied to N 1ref . The value of the signal 88 is subtracted from N 1ref in a subtractor 90, and a signal 92 indicative of the difference is delivered to a subtractor 94 along with the current actual engine low compressor speed signal 44. An error signal 96 indicative of the difference between the actual engine speed N₁ and the desired engine speed represented by the signal 92 is delivered into the control portion 36. When either the engine pressure ratio signal 34 or the EPR schedules within the control portion 36 is unavailable or unreliable the control portion 36 controls the fuel flow via the valve 26 to reduce the error signal 96 to zero. To avoid a "bump" during steady state it is necessary that the trim limiting mechanism of the control system not influence the trim operation when EPR mode failure occurs during steady state engine operation. This will be assured if the value of the signal 68 is always less than the value of the compressor speed which would have been called for by the throttle setting in the EPR mode. The Mach number curves used in the function generator 58 are therefore selected to assure that N E will never be greater than and will preferably be slightly less than the speed called for by the EPR schedule at the time of EPR failure. Although the primary parameter in this embodiment is the engine pressure ratio, it is apparent the present invention may be used with a primary parameter other than EPR (e.g. fan pressure ratio or a weighted fan/engine pressure ratio). In the event of failure of certain parameters required by a control system to operate an engine it is known to calculate a synthesized value of the failed parameter to be used in place of the actual measured parameter. For example, burner pressure is synthesized in commonly owned U.S. patent 4,212,161 to David M. Newirth et al. In the present control system it is desireable to calculate a synthesized value of N₁ (hereinafter referred to as N 1syn ) to be used in the event of unavailability or unreliability of an actual measured value of N₁. As with the sudden switchover to a different control mode, the switch from an actual to a synthesized parameter can result in a "bump". The control logic of Fig. 2 shows a technique for eliminating such a bump. In Fig. 2 the parameter which is synthesized is N₁. In the embodiment of Fig. 2 the parameter which is synthesized is N₁. A synthesized value of N₁ is continuously calculated from the most recent, valid information such that it is always up to date and ready for use upon the actual value of N₁ becoming unavailable or unreliable (i.e., N₁ failure). Referring to the drawing, aircraft Mach number and the high compressor speed corrected to the temperature at the low compressor inlet (N₂C₂) are input to a function generator 200. The function generator 200 produces an output 202 which is an estimated value of the low compressor speed corrected to the temperature at the low compressor inlet (N 1syn C₂). The value of the signal 202 is based upon empirically derived relationships between Mach number and corrected high compressor speed which, in turn, are based upon steady state engine characteristics. In a manner well known in the art, the corrected, estimated low compressor speed 202 and an appropriate multiplier 204 (which would be the same as the multiplier 64 of Fig. 1) are delivered into a multiplier 205 to yield a signal 206 indicative of an estimated or synthesized value of the low rotor speed (N 1syn ). Since the value of N 1syn is based upon empirical relationships between the high and low rotor during steady state engine operation, the signal 206 is passed through a compensator 208 which replaces the dynamic characteristics of the corrected high rotor speed N₂C₂ with that of the output N 1syn C₂ during transient operation. During steady state operation the compensator will have no effect. Compensators of this nature are well known in the art. The N 1syn output signal 99 from the compensator 208, along with the signal 44 indicative of the actual measured value of N₁ are continuously fed to a subtractor 100. A signal 102 representing the difference between them is fed to a switch 104, along with a signal 106 indicating whether or not N₁ has failed (i.e., is unavailable or unreliable). If the signal 106 indicates N₁ is good, the output 108 of the switch 104 will simply be the current calculated difference between N₁ and N 1syn . If the signal 106 indicates N₁ is not good, the output 108 will be the value of the signal 110, which is the last value of the difference between N₁ and N 1syn calculated before the failure signal was received by the switch. In either event, the switch output 108 is sent to a subtractor 112, along with the N 1syn signal 99, and the output signal 114 therefrom is sent to another switch 116, along with a signal 44 indicative of the actual measured value of N₁. The fault signal 106 is delivered to the switch 116. If the signal 106 indicates N₁ is still good, then the N₁ signal 44 is passed therethrough as the switch output signal 118. If N₁ is bad, the signal 118 will have the value of the signal 114, which is an estimated value of N₁. In a preferred embodiment the signal 118 replaces the signal 44 delivered to the subtractor 94 of Fig. 1. In accordance with the present invention, it can be seen that at the instant of N₁ failure (i.e. upon switchover to the use of N 1syn ) the value of the output 114, and thus the value of the signal 118, will be the last good value of N₁ measured essentially at the instant before failure. Thereafter, the value of output 114 will increase and decrease incrementally by an amount equal to the change in the calculated value of N 1syn . Thus, the control produces no bump at the instant of switchover and continues to control the engine smoothly as a function of N 1syn . It should be apparent that the foregoing method for eliminating a bump upon switchover from actual engine speed to synthesized engine speed may be used to eliminate a bump in switching from the use of any measured parameter to the use of a synthesized value of that parameter. Additionally, it should be understood by those skilled in the art that other various changes and omissions in the form and detail of the invention may be made without departing from the spirit and scope thereof.

Claims: 1. In the method of operation of a control system for a gas turbine aircraft engine having a compressor, a burner, and a turbine, wherein the actual value of a first parameter is continuously determined, and wherein a scheduled reference first parameter is compared to the actual value of the first parameter, and the control modulates the fuel flow to the burner to eliminate any difference between the actual and scheduled first parameter, and in the event either the scheduled or actual first parameter cannot be determined, and the first parameter is thereby lost, the control modulates fuel flow to eliminate the difference between a scheduled reference second parameter, which is compressor speed, and the actual value of the compressor speed to eliminate the difference therebetween, and wherein a trim for the scheduled reference compressor speed is computed and applied to the scheduled reference compressor speed upon loss of the first parameter, the trim being computed as the difference between said scheduled reference compressor speed and the actual value of the compressor speed last determined before first parameter mode loss the improvement comprising the steps of: continuously estimating the value of the compressor speed as a function of the scheduled value of the first parameter; continuously calculating the difference between the scheduled reference compressor speed and said estimated value of the compressor speed; and upon first parameter mode loss, limiting the computed trim applied to the scheduled reference compressor speed to the continuously calculated difference last calculated in said preceeding step prior to loss of said first parameter. 2. The improvement according to claim 1 wherein engine pressure ratio is the first parameter, and the step of continuously estimating compressor speed includes estimating compressor speed as a function of the scheduled value of engine pressure ratio. 3. The improvement according to claim 2 wherein said step of continuously estimating compressor speed includes estimating compressor speed as a function of scheduled engine pressure ratio and aircraft Mach number. 4. The improvement according to claim 1 wherein the engine is a twin spool turbofan engine, and said second parameter is low compressor speed. 5. The improvement according to claim 1 wherein said continuously calculated difference is no less than the difference between the scheduled reference compressor speed and the compressor speed which would be called for during the first parameter mode. 6. In a control system for a gas turbine aircraft engine having a compressor, a burner, a turbine and means for continuously measuring the actual value of a first parameter, said control system including means for comparing the actual value of the first parameter to the value of a scheduled reference first parameter; means for generating a first signal when the scheduled or actual value of the first parameter cannot be determined, the first parameter thereby being lost; means for continuously modulating the fuel flow to the burner to eliminate any difference between the actual and scheduled value of the first parameter prior to the first signal being generated; and means for computing a first trim as the difference between the scheduled reference compressor speed and the actual value of the compressor speed last determined before a loss of the first parameter, the improvement comprising: means for determining a trim limit including 1) means for receiving the first signal; 2) means for continuously estimating the value of the compressor speed as a function of the scheduled value of the first parameter; 3) calculating means for continuously calculating a trim limit as the difference between the scheduled reference compressor speed and the estimated value of the compressor speed; and 4) means for generating a second signal upon receipt of the first signal, the second signal being indicative of the trim limit last calculated by said calculating means prior to receiving the first signal; and means for selecting the lower of the first trim and the trim limit, and for applying the selected trim to the scheduled reference compressor speed after the first signal is received; wherein said modulating means includes means for modulating the fuel flow to the burner to eliminate any difference between the actual and trimmed scheduled reference compressor speed after the first signal is received. 7. The improved control system according to claim 6 wherein the first parameter is engine pressure ratio. 8. The improved control system according to claim 7 wherein said estimating means includes means for continuously estimating the value of the compressor speed as a function of both the actual value of engine pressure ratio and aircraft Mach number. 9. The improved control system according to claim 7, wherein said control system is for a twin spool turbofan engine having a low compressor and a high compressor, and the second parameter is low compressor speed.

Inventors: KRUKOSKI, LEON, LAPRAD, RICHARD F.
Applicants: UNITED TECHNOLOGIES CORPORATION
Application year: 1987
Main classification: F02C9
IPC codes: F02C9