The present disclosure relates to an oil system of a gas turbine engine and to a gas turbine engine.

Existing gas turbine engines, respectively so-called turbofan solutions, do not typically incorporate journal bearings. The oil systems are also typically unable to provide an uninterrupted oil supply to the system users, but this is acceptable since roller bearings and gears can tolerate short duration oil interruptions without sustaining damage or overheating. However, gearboxes of gas turbine engines, which are positioned between fan shafts and compressor shafts, incorporate journal bearings that require a constant feed of oil to operate. Failure to provide sufficient oil in the right condition may lead to failure or seizure in the area of a gearbox which would result in a locked fan. If the fan is unable to rotate, this is likely to constitute a hazardous condition to the aircraft, and may e.g. affect the ability to hold a heading.

There are flight manoeuvres which make a sufficient oil supply more difficult. Such a flight manoeuvre is a negative g flight manoeuvre during which the sign of acceleration due to gravity acting on an aircraft changes from positive to negative and remains negative for a certain period of time, e.g. for 5 to 10 seconds. During such a negative g flight manoeuvre the oil in the engine oil tank migrates away from the tank offtake and uncovers the pump which is leading to a drop in oil pressure. This continues until positive g resumes, whereupon oil migrates back to the bottom of the tank and recovers the oil pump. The air which has entered the system is then purged, and normal oil flow resumes.

There is prior art for negative g capable oil systems, and often they

incorporate so-called negative features within the tank, e.g. valves or baffles, to ensure a continuous flow of oil. These systems also rely on the oil being returned to the tank from the oil chambers, and this can be achieved by rotationally dominated flow regimes or through the oil chamber vent system having sufficient capacity to return oil to the tank. Gas turbine engines with gearboxes sometimes do not incorporate a vent system and have large oil flow rates into the gearboxes which do not currently incorporate a scavenge system that would be negative g capable. Hence, introducing features into the tank to ensure that oil is always delivered from the tank could result in the tank being emptied during the negative g flight manoeuvre if little oil is being returned, and therefore, still in an interruption to the oil supply.

If the oil chambers were capable of returning the oil during negative g flight manoeuvres, introducing tank features would still increase complexity, weight and cost.

Even if there may be capability for a few seconds it is still unrealistic to expect this to increase to the full negative g requirement. If journal bearings are frequently insufficiently supplied with oil an early degradation could be the result.

It is the object of the present disclosure to provide an oil system of a gas turbine engine, which ensures a sufficient oil supply for components of a gas turbine engine. Furthermore it is the object of the present disclosure to provide a gas turbine engine of an aircraft which is characterised by low maintenance costs.

This object is achieved through an oil system, and with a gas turbine engine with the features of claim 1 or 11 , respectively.

According to a first aspect, an oil system of a gas turbine engine is provided. The oil system is comprising an oil circuit and at least one further oil circuit. Via the oil circuits at least one journal bearing can be supplied with oil from at least one oil tank. A volume of the oil circuit and a volume of the further oil circuit are correlated with each other in such a way that the journal bearing is supplied with oil during and after completion of a negative g flight manoeuvre via the oil circuits. The oil system according to the present disclosure enables safe engine op- eration during negative g flight manoeuvres or less than zero g flight manoeuvres respectively, leading to decreased cost of ownership.

The oil system is simply using at least two oil paths to deliver oil to a journal bearing. The paths are sized such that the air pockets that enter them during nega- tive g flight manoeuvres are never delivered by each line at the same time. This ensures an uninterruptable oil flow to the journal bearing.

If each oil circuit is comprising at least one pump, the supply with oil through the oil circuits is ensured at least during a positive g operational state of a gas tur bine engine even in case of failure of one of the pumps.

In a simple construction of the oil system the volume of the oil circuit is smaller than the volume of the further oil circuit. This embodiment of the oil system comprises two or more oil circuits of different volumes to feed the journal bearings. The two circuits will have different oil flow characteristics under negative g flight manoeuvres so that the journal bearings continue to receive oil during and after the manoeuvres.

In a further embodiment of the oil system oil is supplied to the journal bearing by both circuits during a positive g operational state of a gas turbine engine. This secures the consistent supply of oil during positive g flight manoeuvres even in case of failure of one of the pumps.

If the oil tank is incorporating offtakes of the oil circuits which are located ei- ther at the same height or at different heights of the oil tank, a consistent supply of oil to the journal bearing is ensured during a positive g operational state as well as during a negative g flight manoeuvre. In this context it is possible to position the oil offtakes of the oil circuits in the oil tank in such a way that the pump of the further oil circuit sucks air from the oil tank later than the pump of the oil circuit during a nega- tive g flight manoeuvre. According to a further aspect of the present disclosure, the volume of the oil circuit and the volume of the further oil circuit are correlated with each other in such a way, that the journal bearing is consistently supplied with oil during a negative g flight manoeuvre, and after the negative g flight manoeuvre via the further oil circuit at least until that moment when the journal bearing is supplied with oil via the oil circuit.

If an additional oil circuit with a separate pump is provided via which oil is suppliable to the journal bearing during a positive g operational state and/or during a negative g flight manoeuvre, a sufficient supply of oil to the journal bearing is ensured in a simple manner.

According to a further aspect of the present disclosure, the volume of the fur ther oil circuit between the pump of the further oil circuit and the journal bearing is greater than the volume of the oil circuit between the pump of the oil circuit and the journal bearing. This embodiment ensures that the journal bearing will be supplied with oil from the further circuit during a negative g flight manoeuvre and after the negative g flight manoeuvre until the oil circuit is purged.

If the flow rates of the pumps correspond to or differ from each other, the oil supply via the oil circuits to the journal bearing is adaptable depending on the re- spective embodiment of the oil system according to the present disclosure.

As noted elsewhere herein, the present disclosure relates to a gas turbine engine. Such a gas turbine engine may include an engine core comprising a turbine, a combustor, a compressor, an oil system according to the present disclosure and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core. The gas turbine engine as described and claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor.

The unit may be designed as a gearbox, especially as a before mentioned epicyclic gear system that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear or from a separate turbine. The core shaft may rigidly connect the turbine and the compressor, such that the turbine and

compressor rotate at the same speed (with the fan rotating at a lower speed).

The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above. In any gas turbine engine as described and claimed herein, a combustor may be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The combustor may be provided upstream of the turbine(s).

The or each compressor (for example the first compressor and second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other.

Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or 0% span position, to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: 0.4, 0.39,

0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31 , 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform.

The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm or 390 cm (around 155 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 250 cm to 300 cm (for example 250 cm to 280 cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 320 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity U _tiP . The work done by the fan blades 13 on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/U _tiP ² , where dH is the enthalpy rise (for example the 1 -D average enthalpy rise) across the fan and U _tiP is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.3, 0.31 , 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in this paragraph being Jkg ¹ K ¹ /(ms ¹ ) ² ). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11 , 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, or 17. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds).

Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg ¹ s, 105 Nkg ¹ s, 100 Nkg ¹ s, 95 Nkg ¹ s, 90 Nkg ¹ s, 85 Nkg ¹ s or 80 Nkg ¹ s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Such engines may be particularly efficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160kN, 170kN, 180kN, 190kN, 200kN, 250kN, 300kN, 350kN, 400kN, 450kN, 500kN, or 550kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 °C (ambient pressure 101.3kPa, temperature 30 °C), with the engine static.

In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: 1400K, 1450K,

1500K, 1550K, 1600K or 1650K. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700K, 1750K, 1800K,

1850K, 1900K, 1950K or 2000K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition. A fan blade and/or aerofoil portion of a fan blade described herein may be manufactured from any suitable material or combination of materials. For example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc. By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a blisk or a bling. Any suitable method may be used to manufacture such a blisk or bling. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding.

The gas turbine engines described and claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN.

The fan of a gas turbine as described and claimed herein may have any desired number of fan blades, for example 16, 18, 20, or 22 fan blades.

As used herein, cruise conditions may mean cruise conditions of an aircraft to which the gas turbine engine is attached. Such cruise conditions may be conventionally defined as the conditions at mid-cruise, for example the conditions experienced by the aircraft and/or engine at the midpoint (in terms of time and/or distance) between top of climb and start of decent.

Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81 , for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions at an altitude that is in the range of from 10000m to

15000m, for example in the range of from 10000m to 12000m, for example in the range of from 10400m to 11600m (around 38000 ft), for example in the range of from 10500m to 11500m, for example in the range of from 10600m to 11400m, for example in the range of from 10700m (around 35000 ft) to 11300m, for example in the range of from 10800m to 11200m, for example in the range of from 10900m to 11100m, for example on the order of 11000m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges. Purely by way of example, the cruise conditions may correspond to: a forward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of -55 °C.

As used anywhere herein,“cruise” or“cruise conditions” may mean the aerodynamic design point. Such an aerodynamic design point (or ADP) may correspond to the conditions (comprising, for example, one or more of the Mach Number, environmental conditions and thrust requirement) for which the fan is designed to operate. This may mean, for example, the conditions at which the fan (or gas turbine engine) is designed to have optimum efficiency.

In use, a gas turbine engine described and claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust.

The skilled person will appreciate that except where mutually exclusive, a feature or parameter described in relation to any one of the above aspects may be applied to any other aspect. Furthermore, except where mutually exclusive, any feature or parameter described herein may be applied to any aspect and/or combined with any other feature or parameter described herein.

Embodiments will now be described by way of example only, with reference to the Figures, in which:

Fig. 1 is a sectional side view of a gas turbine engine;

Fig. 2 is a close up sectional side view of an upstream portion of a gas turbine engine;

Fig. 3 is a partially cut-away view of a gearbox for a gas turbine engine; Fig. 4 an embodiment of an oil system of a gas turbine engine of an aircraft during a positive g operational state of the gas turbine engine;

Fig. 5 the oil system according to Fig. 4 during a negative g flight manoeuvre, wherein air is entering a first oil circuit and a second oil circuit, and oil is still supplied via the first oil circuit and via the second oil circuit of the oil system to a gearbox of the gas turbine engine;

Fig. 6 the oil system according to Fig. 4 during the sustained negative g flight manoeuvre, wherein the first oil circuit is completely vented and oil is supplied only via the second oil circuit to the gearbox;

Fig. 7 the oil system according to Fig. 4 during a positive g flight operation following the negative g flight manoeuvre, whereby the gearbox is only supplied with oil via the second oil circuit, while the first oil circuit is increasingly purged; and

Fig. 8 shows the oil system according to Fig. 4 during a positive g flight operation following the negative g flight manoeuvre, whereby the gearbox is only supplied with oil via the first oil circuit, while the second oil circuit is increasingly purged.

Fig. 1 illustrates a gas turbine engine 10 having a principal rotational axis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23 that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine 10 comprises a core 11 that receives the core airflow A. The engine core 11

comprises, in axial flow series, a low pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, a low pressure turbine 19 and a core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine 10, and defines a bypass duct 22 and a bypass exhaust nozzle 18. The bypass airflow B flows through the bypass duct 22. The fan 23 is attached to and driven by the low pressure turbine 19 via a shaft 26 and an epicyclic gear system 30 which is a planetary gearbox. In use, the core airflow A is accelerated and compressed by the low pressure compressor 14 and directed into the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines 17, 19 before being exhausted through the nozzle 20 to provide some propulsive thrust. The high pressure turbine 17 drives the high pressure compressor 15 by a suitable interconnecting shaft 27. The fan 23 generally provides the majority of the propulsive thrust. The epicyclic gearbox 30 is a reduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shown in Fig. 2. The low pressure turbine 19 (see Fig. 1 ) drives the shaft 26, which is coupled to a sun wheel, or sun gear 28 of the epicyclic gear arrangement 30.

Radially outwardly of the sun gear 28 and intermeshing therewith is a plurality of planet gears 32 that are coupled together by a planet carrier 34. The planet carrier 34 constrains the planet gears 32 to precess around the sun gear 28 in

synchronicity whilst enabling each planet gear 32 to rotate about its own axis. The planet carrier 34 is coupled via linkages 36 to the fan 23 in order to drive its rotation about the engine axis 9. Radially outwardly of the planet gears 32 and intermeshing therewith is an annulus or ring gear 38 that is coupled, via linkages 40, to a stationary supporting structure 24.

Note that the terms“low pressure turbine” and“low pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan 23) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft 26 with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan 23). In some literature, the“low pressure turbine” and“low pressure compressor” referred to herein may alternatively be known as the

“intermediate pressure turbine” and“intermediate pressure compressor”. Where such alternative nomenclature is used, the fan 23 may be referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail in Fig. 3. Each of the sun gear 28, planet gears 32 and ring gear 38 comprise teeth about their periphery to intermesh with the other gears. However, for clarity only

exemplary portions of the teeth are illustrated in Fig. 3. There are four planet gears 32 illustrated, although it will be apparent to the skilled reader that more or fewer planet gears 32 may be provided within the scope of the claimed invention. Practical applications of an epicyclic gearbox 30 generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in Fig. 2 and Fig. 3 is of the planetary type, in that the planet carrier 34 is coupled to an output shaft via linkages 36, with the ring gear 38 fixed. However, any other suitable type of epicyclic gearbox 30 may be used. By way of further example, the epicyclic gearbox 30 may be a star arrangement, in which the planet carrier 34 is held fixed, with the ring (or annulus) gear 38 allowed to rotate. In such an arrangement the fan 23 is driven by the ring gear 38. By way of further alternative example, the gearbox 30 may be a differential gearbox in which the ring gear 38 and the planet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in Fig. 2 and Fig. 3 is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox 30 in the engine 10 and/or for connecting the gearbox 30 to the engine 10. By way of further example, the connections (such as the linkages 36, 40 in the Fig. 2 example) between the gearbox 30 and other parts of the engine 10 (such as the input shaft 26, the output shaft and the fixed structure 24) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of Fig. 2. For example, where the gearbox 30 has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in Fig. 2.

Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations.

Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor).

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in Fig. 1 has a split flow nozzle 20, 22 meaning that the flow through the bypass duct 22 has its own nozzle that is separate to and radially outside the core engine nozzle 20.

However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct 22 and the flow through the core 11 are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. Whilst the described example relates to a turbofan engine, the disclosure may apply, for example, to any type of gas turbine engine, such as an open rotor (in which the fan stage is not surrounded by a nacelle) or turboprop engine, for example.

The geometry of the gas turbine engine 10, and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis 9), a radial direction (in the bottom-to-top direction in Fig. 1 ), and a circumferential direction (perpendicular to the page in the Fig. 1 view). The axial, radial and circumferential directions are mutually perpendicular.

Fig. 4 to Fig. 8 each shows a schematic view of an oil system 42 of the gas turbine engine 10 of an aircraft during different operational states of the aircraft and of the oil system 42. The following description can be read in conjunction with Fig. 4 to Fig. 8, which illustrate the principle concept of the oil system 42 concerning the feed of journal bearings 41 of the gearbox 30.

The oil system 42 comprises two oil circuits 43, 44. Both, the first oil circuit 43 and the second oil circuit 44 are equipped with a pump 45, 46. The journal bearings 41 of the gearbox 30 can be supplied with oil from an oil tank 47 via the oil circuits 43, 44. The oil tank 47 does not comprise any negative features. The volume of the first oil circuit 43 between the oil tank 47 and the journal bearings 41 is smaller than the volume of the second oil circuit 44 between the oil tank 47 and the journal bearings 41. Furthermore, the two circuits 43, 44 have different oil flow

characteristics under negative g flight manoeuvres so that the journal bearings 41 continue to receive oil during and after such a negative g flight manoeuvre which can last for example up to 10 seconds.

In general, the volumes and flow rates of the two oil circuits 43, 44 of the oil system 42 are designed as a function of the maximum permissible duration of a negative g flight manoeuvre. This ensures that the journal bearings 41 are supplied with oil over the entire operating range of the gas turbine engine 10.

During a positive g operational state of the aircraft and of the oil system 42, both oil circuits 43, 44 are delivering oil to the journal bearings 41 at similar flow rates. This operational state and a corresponding filling level 52 of the oil tank 47 of the oil system 42 is illustrated in Fig. 4. Once a negative g flight manoeuvre has begun, both pumps 45, 46 will de-prime and start to suck air. Fig. 5 shows the oil system 42 and a filling level 53 during a negative g flight manoeuvre and at a point in time before the end of a maximum duration of a negative g flight manoeuvre. At this point in time, the air in the oil circuits 43, 44 is compressed by the pumps 45,

46. In addition, a reduced oil flow from pumps 45, 46 to the journal bearings 41 is maintained. In comparison to a positive g operating condition of the aircraft, the oil volume is then pumped via both oil circuits 43, 44 with a reduced flow rate to the journal bearings 41. This results from the air, which is entering the pumps 45, 46 and the delivery lines of the oil circuits 43, 44. In this context it is noted that vented areas of the oil circuits 43, 44 are shown in Fig. 5 to Fig. 8 under the reference numbers 48, 49.

Depending on how long the negative g flight manoeuvre of the aircraft lasts, the first oil circuit 43 with the smaller volume may be emptied before the second oil circuit 44 with the larger volume. Then, oil is delivered to the journal bearings 41 via the second oil circuit 44 only whilst air fills the delivery lines. This operational state of the oil system 42 is shown in Fig. 6. The volume of the second oil circuit 44 is sized so as to always deliver oil for the length of the negative g flight manoeuvre with margin.

Once the gas turbine engine 10 and the oil in the oil tank 47 are experiencing positive g, the first oil circuit 43 re-primes and purges its air before the large volume circuit 44 has finished delivering oil and started to purge its air. Fig. 7 shows an operational state of the oil system 42 during a positive g operational state of the aircraft following the negative g flight manoeuvre. During this operational state, the oil is supplied to the journal bearings 41 only via the second oil circuit 44 and both oil circuits 43, 44 are purged with oil from offtakes 50, 51 of the oil tank 47. The second oil circuit 44 provides oil for the journal bearings 41 at least until that point in time when the first oil circuit 43 is completely purged with oil and the journal bearings 41 are supplied with oil also via the second oil circuit 44.

Then, the journal bearings 41 are only supplied with oil via the first oil circuit 43 (as shown in Fig. 8) until the second oil circuit 44 is completely purged and the journal bearings 41 are also supplied with oil via the second oil circuit 44. The oil system according to the present disclosure may comprise at least two separate pumps, drawing upon oil from one reservoir or from different reservoirs. The volume of the pipework of the oil circuits between the pumps and the journal bearings needs to be sufficiently different between the two delivery paths to ensure that with their respective flow rates, neither interrupts flow to the journal bearings at a coincident time.

Such an embodied oil system will ensure during a sustained negative g flight manoeuvre that the journal bearings will receive a continuous oil flow. It will be a reduced flow rate compared to nominal but this will be sufficient to avoid

degradation of the journal bearings.

Parts list principal rotational axis

engine

core

air intake

low-pressure compressor

high-pressure compressor

combustion equipment

high-pressure turbine

bypass exhaust nozzle

low-pressure turbine

core exhaust nozzle

nacelle

bypass duct

propulsive fan

stationary supporting structure shaft

interconnecting shaft

sun gear

epicyclic gear system

planet gears

planet carrier

linkage

ring gear

linkage

journal bearing

oil system

, 44 oil circuit

, 46 pump

oil tank

, 49 vented area 50, 51 offtake

52, 53 filling level

A core airflow

B bypass airflow