GB1597173A

GB1597173A - Device for optimum control of aero gas turbine engines

Info

Publication number: GB1597173A
Application number: GB51154/77A
Authority: GB
Original assignee: MTU Motoren und Turbinen Union Muenchen GmbH
Current assignee: MTU Aero Engines AG
Priority date: 1977-04-05
Filing date: 1977-12-08
Publication date: 1981-09-03
Also published as: DE2715300B2; FR2386686A1; IT1101911B; DE2715300C3; IT7847587A0; FR2386686B1; DE2715300A1

Description

(54) DEVICE FOR OPTIMUM CONTROL OF AERO GAS TURBINE ENGINES (71) We, MOTOREN-UND TURBINEN-UNION MUNCHEN GmbH, a German Limited Liability Company, of 8 Munchen 50, Postfach 50 06 40, Germany, do hereby declare the invention for which we pray that a patent may be granted to us, and the method by which it is to be performed, to be particularly described in and by the following statement: The invention relates to a device for optimum control of aero gas turbine engines, in which, in order to achieve optimum control of several relevant engine variables, additional information is used in the form of functions of the intake air condition and the actual engine operating mode at a given moment.

The engine parameters for which optimum control is to be provided may, for instance, be the fuel supply, nozzle area, angular setting of vanes on the stator of the turbine and/or compressor. and the setting of air bleed valves, etc.

Normally. modern aero gas turbine engines incorporate aerodynamically highly loaded components with a relatively narrow operating range. In order to ensure satisfactory.

operation of such engines, further parameters are frequently used in addition to the natural parameters in the form of fuel flow to the combustion chamber and to the afterburner (if there is one). Such additional parameters include predominantly a variable exhaust gas nozzle, variable stators in some compressor stages, as well as bleed valves at one or more locations in the engine.

These parameters can be controlled in a number of quite different ways. In many cases, a simple control is used, e.g., a schedule of the compressor stator vanes as a direct function of corrected speed, while bleed valves can be opened below a certain speed or engine pressure. Adjustment of the nozzle is frequently effected by control methods by means of which certain parameters, such as the turbine inlet temperature or the compressor working line, are controlled to specified values. The design of all these control mechanisms has to be such that satisfactory operation is ensured even during critical flight manoevres and/or engine operating conditions.

Since advanced aircraft propulsion systems, particularly in the military sector, have frequently to cater for the most varying flight conditions and manoeuvres, and since the requirements for rapid changes of engine operation conditions have grown considerably, the control concepts known up to now often require rather big compromises under many non-critical operating conditions, only in order to avoid compressor surging, overtemperature and similar unacceptable problems in just a few critical cases.

To illustrate this briefly, the following example is given. Normally, the selection of the nozzle area for optimum fuel consumption in operating a bypass engine without reheat depends upon the efficiency characteristics in the low-pressure compressor as well as on the afterbody drag of the aircraft concerned. It is quite possible that an extreme turn or slow flight with large incidence leads to such an unfavourable pressure profile at the intake of the first compressor that surging would occur in operation on the fuel-consumption-wise optimum LP-compressor working line. Solely to cater for such relatively rare cases, a larger nozzle area is necessary, which results in increased specific fuel consumption and reduced range of the aircraft.

In the light of the problems described above, one object of the present invention is to devise a control device for aero gas turbine engines, by means of which the compromises made normally at the expense of overall economy can largely be avoided with relatively small effort in terms of control hardware.

In addition, it is intended to make the available variables of an aero gas turbine engine flexibly adaptable in an optimum way to actual requirements existing at any given time.

For this purpose, information from two different sources is required; these are; - condition of the air at the engine intake - operating condition selected (steady-state and non-steady-state) by the pilot.

It is assumed, in the first place, that the control systems for the basic engine and the afterburner (if there is one) are optimized for ideal steady-state operation without or with constant pressure profile at the engine inlet face, in order to achieve the maximum possible thrust and the lowest possible specific consumption. To this end, numerous differing methods of control are known and in use, which are, however, not the subject of the present invention.

To solve the problems outlined above we propose a device for optimum control of an aero gas turbine engine wherein the throttle lever controls the engine operating mode and determines, via an engine control unit, the setting of various engine parameters such as main and afterburner fuel flow and the exit area of a variable nozzle at the jet pipe exit, the device comprising a first electronic data processing unit for calculating the engine operating mode set by the pilot from data representing the position and the rate of change in position of the throttle lever, a second electronic data processing unit for calculating the Reynolds No. at the engine intake and the type and magnitude of the inlet pressure profile from data measured during flight (e.g. overall air temperature and pressure, Mach No., drift angle of the aircraft or unambiguous functions of these parameters), and a third electronic data processing unit for generating, in response to the data representing the set operating mode and the flight condition derived respectively by the first and second processing units, control signals, corresponding to the optimum setting of the engine parameters which override or trim the standard control functions of the engine control unit.The device is responsive to information received from the two above-mentioned sources and able to recognize critical conditions for the engine and to influence the standard control functions of the engine by means of trims in such a way that, through appropriate changes in one or more of the engine variables, the engine is not endangered for the duration of the critical condition. Of course, such measures are also, in most cases, at the expense of optimum thrust and fuel consumption; since, however, such critical conditions are generally of short duration, the most economic solution is thus achieved, because during the subsequent transition to normal operating conditions the engine variables can also be re-set to optimum values of consumption and/or thrust by the basic control system.

For the determination of the air condition at the engine intake, the following data are generally required: - overall air temperature - overall air pressure - flight Mach No.

- side-slip angle of the aircraft - position of the variable geometry in the air intake.

From these data, the information require for operation of the device according to this invention can be derived: - level of Reynold's No. for engine components - type of intake pressure profile (e.g., radially, circumferentially or combination of both) - magnitude of the pressure profile.

Type and magnitude of the pressure profile as a function of the above data can be obtained bv means of engine test runs in an altitude test facility with the intake installed, or as results of less expensive model tests.

Moreover. the effects of Revnold's No. changes, as well as of the various types of intake pressure profiles which affect the performance characteristics of the individual engine components. are known from tests with complete engines or from separate component test runs. From them. the necessary adjustments to the engine parameters can be determined relatively easily.

There is. however, an important second source of information defined by the requirements the pilot makes on the engine. This information can be used as an additional control of the engine parameters. If the pilot, for example, runs the engine in the throttled-back operating range, the parameters are controlled in such a way that the engine is trimmed optimallv with regard to fuel consumption - under certain circumstances at the expense of maximum thrust, which is not required in this condition anyway.

On the other hand. the specific consumption of the basic engine will play a minor role when maximum thrust is selected, particularly with afterburner operation, since the engine is generally not run in this condition for very long. Similar examples can easily be given for rapid load changes required by the pilot. The information defining the requirements of the pilot at a given moment is - for the purposes of the present invention - obtained above all from the position a or the change of position a of the throttle lever.The table shows the logic upon which this concept is based: No. operating condition signal from requirement selected by the pilot throttle lever for engine 1 idle a=a idle lowest thrust 2 cruise (xidle < o! < a lowest fuel full load, consumption afterburner unlit 3 full load without a=a maximum thrust afterburner full load, without after afterburner burner unlit 4 maximum afterburning a=a maximum thrust full load, with after- with after- burner burning 5 acceleration without & gt;;K rapid and afterburning a idle > (x > a reliable full load, acceleration afterburner without com unlit pressor or dN/dt > K2 surging 6 deceleration. after- & rapid and reliable burner unlit a idle < α < α deceleration with full load, out compressor afterburner surging or flame unlit out or too rapid or N/dt < K change of the turbine inlet temperature 7 running up of the a > a rapid and reliable afterburner a min AB > α< αmax AB acceleration, correct afterburner fuel metering with dANozzIe/dt > K4 out compressor surging 8 running down of o < x rapid and reliable the afterburner a min AB < o < oLmax AB deceleration, correct afterburner fuel metering with dANozzleIdt < K5 out compressor surging In the above table K. to K5 = given constants a = position of the throttle lever & = shifting speed of the throttle lever Nozzle = nozzle area N = rotational speed dN/dt = rotational acceleration AB = afterburner The eight typical cases listed should be considered as examples only.The concept can be extended to include operations such as starting and shutdown of the engine, actuation of the thrust reverser, firing of weapons, etc.

It is important to note that the required optimum setting of the engine parameters is a function of the intake data as well as of the information derived from the throttle lever.

One embodiment of the present invention will now be described by way of example with reference to the accompanying drawings which is a block diagram of a device for optimum control of an aero gas turbine engine.

In a data processing unit 2, the Reynold's No. level and the type and magnitude of the intake pressure profile is calculated from the data measured continuously during flight and from the setting of the variable intake geometry, all of which is available from the aircraft flight recorder 1.

Together with the operating mode (calculated in data processing unit 4) of the aero gas turbine engine 8 selected by the pilot by means of the throttle lever (unit 3), the information obtained from units 1,2 and 3,4 is used to determine the optimum setting of the engine parameters in data processing unit 5 producing output signals for overriding the basic engine control system (engine and afterburner control unit 6).

The throttle lever position 3 input is also connected by line 9 directly to the engine and afterburner control unit 6.

The aero gas turbine engine 8 shown schematically may be, for instance, a three-shaft bypass engine with afterburner, and a variable nozzle 10.

The present invention can also be applied to other types of aero gas turbine engines, for example, to a multi-shaft, multi-stream engine or a pure straight jet engine.

The engine parameters of the gas turbine jet engine 8 shown include, apart from the engine fuel flow to the injection nozzles of the main combustion chamber (signal 11 from unit 6), the afterburner fuel flow (signal 12 from unit 6). In addition, the engine parameters include. for example, the angular setting of the compressor stator guide vanes and the setting of the compressor first and second bleed valves, the signal flow for the control of the last-mentioned three engine variables being identified by item Nos. 13, 14 and 15 respectively. A further engine variable is the nozzle area of the variable nozzle 10 (signal flow 16 from unit 6). Signal flow 17 from the engine to unit 6 identifies all engine parameters (speeds, pressures, temperatures) which are required for engine and afterburner control.

The engine parameters may, in addition, include bleed valves at several stations of the thermodynamic cycle, and the angular setting of vanes on the turbine stator; there is, however, no indication of this on the accompanying drawing.

The precise logic to be used in unit 5 depends largely upon the specific engine type with its particular aerodynamic characteristics of the individual components and will certainly have to be varied from case to case. Therefore, the present invention is only intended to state, as a formality, the dependencies in generally valid form. It is pre-supposed that the bleed valves are valves which are normally closed, whereas the variable stators and the nozzle area are first of all adjusted in accordance with a definite schedule by the engine and afterburner control system, with the position being trimmed by a certain amount through the logic in unit 5.

Position of bleed valve I = fl (Reynold's No., type of pressure profile, magnitude of pressure profile, engine operating condition) Position of bleed valve II = f2 (Reynold's No., type of pressure profile, magnitube of pressure profile, engine operating condition) Aa stators = fX (Reynold's No., type of pressure profile, magnitude of pressure profile, engine operating condition) AA nozzle area = f4 (Reynold's No., type of pressure profile, magnitude of pressure profile, engine operating condition) The above specification can be continued at will for other engine parameters. Data processing units 2 and 4 are electronic computing devices, and may be either analogue or, preferably, digital devices.

WHAT WE CLAIM IS: 1. A device for optimum control of an aero gas turbine engine wherein the throttle lever controls the engine operating mode and determines, via an engine control unit, the setting of various engine parameters, the device comprising a first electronic data processing unit for calculating the engine operating mode set by the pilot from the data representing the position and the rate of change in position of the throttle lever, a second electronic data processing unit for calculating the Reynolds No. at the engine intake and the type and magnitude of the inlet pressure profile from data measured during flight and a third electronic data processing unit for generating. in response to the data representing the set operating mode and the flight condition derived respectively by the first and second

**WARNING** end of DESC field may overlap start of CLMS **.

Claims

**WARNING** start of CLMS field may overlap end of DESC **. from the setting of the variable intake geometry, all of which is available from the aircraft flight recorder 1. Together with the operating mode (calculated in data processing unit 4) of the aero gas turbine engine 8 selected by the pilot by means of the throttle lever (unit 3), the information obtained from units 1,2 and 3,4 is used to determine the optimum setting of the engine parameters in data processing unit 5 producing output signals for overriding the basic engine control system (engine and afterburner control unit 6). The throttle lever position 3 input is also connected by line 9 directly to the engine and afterburner control unit 6. The aero gas turbine engine 8 shown schematically may be, for instance, a three-shaft bypass engine with afterburner, and a variable nozzle 10. The present invention can also be applied to other types of aero gas turbine engines, for example, to a multi-shaft, multi-stream engine or a pure straight jet engine. The engine parameters of the gas turbine jet engine 8 shown include, apart from the engine fuel flow to the injection nozzles of the main combustion chamber (signal 11 from unit 6), the afterburner fuel flow (signal 12 from unit 6). In addition, the engine parameters include. for example, the angular setting of the compressor stator guide vanes and the setting of the compressor first and second bleed valves, the signal flow for the control of the last-mentioned three engine variables being identified by item Nos. 13, 14 and 15 respectively. A further engine variable is the nozzle area of the variable nozzle 10 (signal flow 16 from unit 6). Signal flow 17 from the engine to unit 6 identifies all engine parameters (speeds, pressures, temperatures) which are required for engine and afterburner control. The engine parameters may, in addition, include bleed valves at several stations of the thermodynamic cycle, and the angular setting of vanes on the turbine stator; there is, however, no indication of this on the accompanying drawing. The precise logic to be used in unit 5 depends largely upon the specific engine type with its particular aerodynamic characteristics of the individual components and will certainly have to be varied from case to case. Therefore, the present invention is only intended to state, as a formality, the dependencies in generally valid form. It is pre-supposed that the bleed valves are valves which are normally closed, whereas the variable stators and the nozzle area are first of all adjusted in accordance with a definite schedule by the engine and afterburner control system, with the position being trimmed by a certain amount through the logic in unit 5. Position of bleed valve I = fl (Reynold's No., type of pressure profile, magnitude of pressure profile, engine operating condition) Position of bleed valve II = f2 (Reynold's No., type of pressure profile, magnitube of pressure profile, engine operating condition) Aa stators = fX (Reynold's No., type of pressure profile, magnitude of pressure profile, engine operating condition) AA nozzle area = f4 (Reynold's No., type of pressure profile, magnitude of pressure profile, engine operating condition) The above specification can be continued at will for other engine parameters. Data processing units 2 and 4 are electronic computing devices, and may be either analogue or, preferably, digital devices. WHAT WE CLAIM IS:

1. A device for optimum control of an aero gas turbine engine wherein the throttle lever controls the engine operating mode and determines, via an engine control unit, the setting of various engine parameters, the device comprising a first electronic data processing unit for calculating the engine operating mode set by the pilot from the data representing the position and the rate of change in position of the throttle lever, a second electronic data processing unit for calculating the Reynolds No. at the engine intake and the type and magnitude of the inlet pressure profile from data measured during flight and a third electronic data processing unit for generating. in response to the data representing the set operating mode and the flight condition derived respectively by the first and second

processing units, control signals corresponding to the optimum setting of the engine parameters which override or trim the standard control functions of the engine control unit.

2. A device according to claim 1, which is mounted in an aircraft and connected to the aircraft flight recorder whereby the data representing the flight condition is obtained.

3. A device according to claim 1 or claim 2 wherein the engine parameters include, the engine fuel flow and the afterburner fuel flow.

4. A device according to any one of claims 1 to 3, wherein the engine parameters include the exit areas of one or more variable nozzles at the jet pipe exit.

5. A device according to any one of claims 1 to 4, wherein the engine parameters also include the angular setting of vanes on the stator of the compressor.

6. A device according to any one of claims 1 to 5, wherein the engine parameters include the angular setting of vanes on the stator of the turbine.

7. A device according to claims 1 to 6, wherein the engine parameters include the settings of air bleed valves at various stages in the thermodynamic cycle.

8. A device according to any one of claims 1 to 7, wherein the electronic data processing units are analogue units.

9. A device according to any one of claims 1 to 7 wherein the electronic data processing units are digital units.

10. A device for the control of an aero gas turbine engine constructed and arranged substantially as hereinbefore described with reference to and as illustrated in the accompanying drawing.

11. A method of controlling an aero gas turbine engine comprising using a device which calculates data representing the Reynold's Number at the engine air intake and the type and magnitude of the intake pressure profile and the mode of operation of the engine and automatically optimises the engine operating parameters by overriding or trimming the engine control system.

12. A method of controlling an aero gas turbine engine substantially as hereinbefore described.