GB2197391A - Fuel system for gas turbine aero engine - Google Patents

Abstract

A fuel pump (26) is provided with a pump spill return loop for recirculating through the pump any fuel already pumped but not passing through the metering unit (28). Such repumping of fuel can cause it to become excessively hot under some engine operating conditions and to avoid this the control unit (32') is made to produce both a normal fuel demand signal (30) in accordance with a normal control schedule and an alternative fuel demand signal (F), the latter being produced in accordance with an alternative control schedule (54) as a function of fuel temperature (t). The alternative control schedule is such that at fuel temperatures higher than a limiting value (tL), the alternative fuel demand signal is the higher of the two, and is selected to control the metering unit (28) by a "highest wins" logic (56), thereby reducing the heat input to the fuel by reducing the amount of fuel being recirculated and repumped. <IMAGE>

Description

SPECIFICATION Fuel control system for gas turbine aero engine The present invention relates to a fuel system for a gas turbine aero engine, in particular one capable of minimising heat input to the fuel as it passes through the fuel system during certain operational conditions of the aircraft.

Various factors in recent fuel system designs for large turbofan aero engines have had the effect of putting a large amount of heat into the fuel as it flows through the system.

When certain extreme operational conditions coincide with times when the aircraft's fuel tank temperature is high, excessive fuel temperatures can result. High fuel temperatures are undesirable because of the danger of excessive production of fuel vapour and air bubbles during pumping and throttling operations, which themselves put additional heat into the fuel during its passage from the fuel tanks to the combustor of the engine. It should be noted that high fuel temperatures in the fuel tanks are due to high air temperatures at ground level combined with the effect of short journey lengths, which reduce the amount of time which the aircraft spends at the colder high altitudes and hence reduces the amount of heat which the aircraft can lose during flight.

A typical fuel system for a large turbofan aero engine incorporates a high pressure fuel pump. This delivers fuel to the engine through a variable orifice metering valve which is used to throttle the discharge flow of the pump and thereby vary the fuel flow rate to the engine.

The position of the valve is controlled by a control unit which accepts various input signals representative of conditions in the engine, applies control laws to them (as known in the industry), and produces a valve position control signal, which can be understood as a fuel demand signal. Because the pump's fuel pumping capacity at engine idling speeds far exceeds the demand, the pump requires means for recirculating any excess of pumped fuel from the pump outlet back to the pump inlet.Consequently the pump is provided with a spillway return passage connecting the pump outlet to the pump inlet, and this is controlled by a pressure responsive fuel valve so that the lower the fuel demand, the greater will be the proportion of the pump throughput which is recirculated through the spillway and the greater will be the amount of heat put into a given amount of fuel by the repeated repumping and recirculation.

One important factor in fuel system design which has contributed to the problem of fuel overheating is the desire to save fuel by the use of very low fuel flow rates when the engine is idling, either at ground idle speed, or particularly at flight idle speed. It will be- realised that low fuel flow rates to the engine may raise already high fuel temperatures to unacceptable levels because of the heat input to the fuel due to the aforementioned recirculation and repumping.

The present invention minimises the above problem of excessive fuel temperature by modifying the mode of operation of the fuel control system when this is appropriate.

According to the present invention, a fuel system for a gas turbine aeroengine includes pump means for supplying fuel to the engine, control means for producing a normal fuel demand signal, and fuel metering means for metering fuel flow to the engine in accordance with the value of the fuel demand signal, the pump means being provided with fuel recirculation means for recirculating through the pump means any fuel which has already been pumped thereby but which does not pass through the fuel metering means, the fuel system further including sensor means for producing a signal representative of fuel temperature, said control means being adapted to receive the fuel temperature signal and in response thereto to supply an alternative fuel demand signal to the metering means only when said fuel temperature signal exceeds a predetermined value, said supplied alternative fuel demand signal being higher than said -nor- mal fuel demand signal, thereby to reduce the heat input to the fuel by reducing the amount of fuel being recirculated through the pump means.

Preferably, the control means includes means for continuously synthesising said alternative fuel demand signal as a function of the fuel temperature, the arrangement being such that the alternative fuel demand signal is higher than the normal fuel demand signal only when the fuel temperature signal exceeds the predetermined value.

Preferably, the control means further includes means for comparing the values of the alternative and normal fuel demand signals and selecting the alternative fuel demand signal for supply to the metering means in preference to the normal fuel demand signal whenever the alternative fuel deniand signal is higher than the normal fuel demand signal.

Exemplary embodiments of the invention will now be described with reference to the accompanying drawings, in which: Figure 1 is a simplified block diagram of a known fuel system for a gas turbine aero engine; and Figure 2 is a block diagram illustrating how the invention can be applied to modify a fuel system like that in Figure 1.

Referring to Figure 1, there is shown in simplified diagrammatic form a fuel system 10 of known type. The fuel system 10 supplies fuel to the gas turbine aero engine (not shown) and also controls the exact amount of fuel being supplied. In the fuel system 10 a first fuel pump 12 draws fuel 14 from the fuel tanks of the aircraft and supplies it at low pressure to the rest of the system. After low pressure pumping in pump 1 2 the fuel passes through a heat exchanger 16, in which it exchanges heat with lubricating oil 18 which circulates in the engine's oil system (not shown).

Normally the fuel is colder than the oil circulating in the engine oil system and therefore the fuel is used to cool the oil. Consequently the heat exchanger 16 is known as a fuel cooled oil cooler, and it also ensures that maximum heat is put into the fuel so as to avoid icing of the fuel filter 20 which is interposed between the oil cooler 16 and the following parts of the fuel system; icing of the filter would of course only occur when the fuel in the aircraft fuel tanks is at or below 0 C. On the other hand, if the fuel from the fuel tanks is already very warm before passing through the fuel system 10, it is sometimes necessary for the oil 18 to bypass the fuel cooled oil cooler 16 so that excessive fuel temperatures in the fuel system are avoided.The danger here is that the fuel will become hot enough to vaporise significantly, causing severe cavitation during subsequent pumping. In such circumstances, heat input from the oil to the fuel is avoided by activating a bypass valve 24 associated with the oil cooler 16. When activated the bypass valve 24 diverts the oil 18 so that instead of passing through the oil cooler 16 it passes through the bypass loop 22. No means for activating the bypass valve 24 are shown, but this could be either manual activation from the flight station upon receiving a fuel temperature warning signal, or automatically by means of a control unit attached to the valve 24, which would activate the valve when the fuel temperature exceeded a predetermined value.

After filtering, the fuel 14 is pumped up to a high pressure by a second fuel pump 26, the output of which is throttled by a fuel metering unit 28. The latter item controls the output of the pump 26 in accordance with a fuel demand signal 30 which is produced by an electronic control unit 32 to which the fuel metering unit 28 is connected. The control unit 32, which may be digital or analogue in nature, receives a number of monitor signals S from various speed, temperature and pressure sensors in the engine, these signals representing various conditions in the engine. It then applies certain predetermined control laws to the monitor signals S and thereby produces the fuel demand signal 30.

Finally the output of the fuel metering unit 28 is passed to the engine through a shutoff valve 34 whose purpose is to completely prevent fuel flow to the engine when the engine is shut down or before it is started, this being achieved by an electrical actuation signal 36 from the flight station.

It will have been previously noticed that the high pressure fuel pump 26 is provided with a spill return valve 38; this is responsive to the pressure drop between the points A and B across the fuel metering unit 28 so that whenever the output of the pump 26 is greater than that being allowed through to the engine by the fuel metering unit 28, the surplus fuel is recirculated to the inlet side of the pump 26 through the spill return loop 40. The spill return valve 38 is of course actuated by the pressure difference between points A and B, being directly connected to the output of the high pressure pump 26 through the spill return loop 40 and also receiving a pressure signal from B on a hydraulic line 42.

The fuel system 10 may be thought of as including a fuel control subsystem comprising the fuel metering unit 28 and the control unit 32, together with the high pressure fuel pump 26 and the shut off valve 36. This fuel control sub system is required to give the pilot of the aircraft control of the thrust of the gas turbine aero engine and to ensure smooth acceleration and deceleration under all conditions by controlling the inputs of fuel to the engine from the fuel system 10. The pilot exerts control over the thrust of the engine by means of control input T to the control unit 32. Control input T is dependant upon the thrust lever setting selected by the pilot in the flight station and together with the monitor signals S determines the value of the fuel demand signals 30 in accordance with the control laws which are programmed or otherwise built in to the control unit 32.

Referring now to Figure 2, it is shown how the fuel control subsystem of Figure 1 may be modified in accordance with the invention. The subsystem is modified in two basic respects: firstly, a temperature sensor 50 is incorporated in the fuel supply line either just before the filter 20, or just after it (latter position shown by dotted lines), and gives a continuous signal t which is a measure of the fuel temperature; secondly, the control unit 32' is different from control unit 32 (Figure 1) in that although the former still incorporates the same engine control laws 52 as the latter in order to give a basic or "normal" fuel demand signal 30, it also incorporates an alternative fuel flow control law 54 which operates on the fuel temperature signal t and produces an alternative "synthetic" fuel demand signal F as a function of t, both of the signals F and 30 being input to a "highest wins" logic 56 which selects whichever of the two signals is the higher and outputs it from the control unit 32' as the actual fuel demand signal 30'. The rest of the fuel control subsystem is identical to that described in connection with Figure 1.

The above-mentioned modifications are based on the premise that the fuel temperature, as sensed by the sensor 50, only becomes undesirably high at extreme (hence, rare) operational conditions, corresponding to the engine being at flight idling speed at a great altitude (e.g. at the beginning of a gradual descent), with high air temperatures on the ground and resulting high aircraft fuel tank temperatures. When fuel comes into the system from tanks at normal temperatures, the high pressure pump outlet temperature is not unacceptably high and the Figure 1 arrangement would be satisfactory, remembering that the fuel cooled oil cooler 16 can be bypassed by the oil 18 to reduce heat input to the fuel if necessary.Consequently, the control law 54, which within limits can be represented generally as a fuel flow characteristic F = f(t) + c, provides a minimum below which the alternative fuel demand signal F cannot fall. In this example, the characteristic gives F a steady level c for any value of fuel temperature t below a limiting value t,. On those rare occasions when t rises above the limiting value tL, it is arranged that control law 54 produces a value of F which is greater than the value of signal 30 being produced by the basic engine control laws, hence F becomes the output 30' of control unit 32'. Since the fuel metering unit 28 is thus being provided with a higher fuel demand signal, more fuel is passed to the engine, resulting in less being recirculated through the high pressure fuel pump 26 due to the operation of spill return valve 38.This reduces the heat input to the fuel because the amount of fuel subjected to repumping at high pressure is reduced. Hence the fuel temperature at the output of the high pressure pump stays within acceptable limits.

The effect of the above modifications to the fuel control subsystem is to "reset" the engine control laws in the control unit 32' whenever the fuel temperature limit t, is reached.

Of course, the relevant engine control laws are those controlling engine idling speed, since these are the only ones in unseat the extreme operating conditions for which the invention operates.

It will be apparent to the informed reader, therefore, that the invention does not normally affect day-to-day aircraft or engine operation.

Under the extreme conditions when the control laws are reset, there will be a higher descent thrust and greater fuel consumption, but these disadvantages are judged acceptable in view of the benefits 6f the invention.

Claims (4)

1. A fuel system for a gas turbine aeroengine, including pump means for supplying fuel to the engine, control means for producing a normal fuel demand signal, and fuel metering means for metering fuel flow to the engine in accordance with the value of the fuel demand signal, the pump means being provided with fuel recirculation means for recirculating through the pump means any fuel which has already been pumped thereby but which does not pass through the fuel metering means, the fuel system further including sensor means for producing a signal representative of fuel temperature, said control means being adapted to receive the fuel temperature signal and in response thereto to supply an alternative fuel demand signal to the metering means only when said fuel temperature signal exceeds a predetermined value, said supplied alternative fuel demand signal being higher than said normal fuel demand signal thereby to reduce the heat input to the fuel by reducing the amount of fuel being recirculated through the pump means.

2. A fuel system according to claim 1 in which the control means includes means for continuously synthesising said alternative fuel demand signal as a function of the fuel temperature, the arrangement being such that the alternative fuel demand signal is higher than the normal fuel demand signal only when the fuel temperature signal exceeds the predetermined value.

3. A fuel system according to claim 2 in which the control means includes means for comparing the values of the alternative and normal fuel demand signals and selecting the alternative fuel demand signal for supply to the metering means in preference to the normal fuel demand signal whenever the alternative fuel demand signal is higher than the normal fuel demand signal.

4. A fuel system substantially as described in this specification with reference to and as illustrated by Figure 2 of the accompanying drawings.