Gas turbine engine system

An aircraft mounted gas turbine engine system comprises a three shaft propulsive gas turbine engine having three electrical generators; each one respectively associated with one shaft of the engine. Two of the electrical generators are additionally configured so as to function as electric motors so as to facilitate power transfer between the engine shafts. One of the electric generators provides the primary source of electrical power for the aircraft, while the remaining electrical generators provide electrical power for the engine and back-up power for the aircraft.

THE FIELD OF THE INVENTION 
This invention relates to aircraft mounted gas turbine engine systems and 
in particular to the manner in which such systems generate and utilise 
electrical and other power. 
BACKGROUND OF THE INVENTION 
An aircraft, and the gas turbine engines that power it, usually include 
numerous units that require electrical, hydraulic or pneumatic power in 
order to function. Conventionally, the power for such units is derived 
from one of three sources depending upon the operating status of the 
aircraft. If the aircraft is stationary and it is necessary for certain 
electrical support systems to be operational, a ground-based electrical 
generator is coupled to the aircraft. Alternatively, the electrical power 
may be derived from an aircraft mounted generator which is powered by an 
auxiliary power unit mounted on the aircraft. Such an auxiliary power unit 
is commonly in the form of a small gas turbine engine mounted in the tail 
region of the aircraft. 
If the aircraft is airborne, power for the units is derived from the 
engines which power the aircraft. Typically, each engine is provided with 
an auxiliary shaft that transmits power from one of the main shafts of the 
engine to a gearbox mounted on the engine external casing. The gearbox in 
turn drives various electrical generators and hydraulic pumps. A single 
electrical generator may be driven by the gearbox. However in the event of 
the failure of that generator, there would, of course, be an interruption 
in the supply of electrical power to the engine and aircraft. In a 
multi-engined aircraft such an interruption may be acceptable for a short 
time period since electrical power is still likely to be available from 
the remaining engine or engines. However, if the aircraft is required to 
fly in regions in which a diversionary airfield is some distance away, it 
is usually necessary for the gearbox to drive an additional electrical 
generator for back-up purposes. 
Such duplication of electrical power generators leads to difficulties in 
conveniently positioning them relative to the gearbox and also to complex 
and heavy modifications to the gearbox to enable it to drive both 
generators. 
SUMMARY OF THE INVENTION 
It is a general object of the present invention to provide an aircraft 
mounted gas turbine engine system in which such drawbacks are 
substantially avoided. 
According to the present invention, an aircraft mounted gas turbine engine 
system comprises a propulsive gas turbine engine having a core unit that 
includes compressor and turbine portions and a propulsive fan driven by 
said core unit, said engine including a plurality of independent shafts 
drivingly connecting said turbine portions to said compressor portions and 
to said fan, each of said shafts independently and directly driving an 
electrical generator, one of said electrical generators operationally 
constituting the primary source of electrical power for the aircraft 
carrying said engine. 
Since each electrical generator is independently and directly driven by one 
of the main engine shafts, there is no longer any requirement for the main 
engine gearbox to be capable of driving two generators and so it may be 
suitably down-sized or even eliminated. Moreover, the generators may be 
positioned in convenient positions which are not necessarily close to the 
engine gearbox if such a gearbox is fitted. 
If the generators are not associated with the main engine gearbox, it then 
becomes possible to mount them within the core unit of the gas turbine 
engine. Moreover, if the electrical generators are of the type that also 
function as electric motors, it is possible to transfer power between the 
engine shafts in the manner described in our PCT patent application Ser. 
No. WO 95/02120. It also permits the use of at least one of the electric 
motors for engine starting purposes. 
We prefer that in addition to each of the electrical generators performing 
the function of an electric motor, they are configured so that they also 
function as electromagnetic bearings. This brings important advantages in 
terms of simplifying the overall structure of the engine and providing the 
means for greater control over engine operation. 
If electromagnetic bearings are employed in the engine, we prefer that the 
auxiliary power unit of the aircraft on which the engine is mounted is 
integrated into the gas turbine engine system. By doing so, the auxiliary 
power unit can be used for providing the electrical power for 
electromagnetic bearing levitation, for instance, prior to the operation 
of the gas turbine engine.

DETAILED DESCRIPTION OF THE INVENTION 
With reference to FIG. 1, a ducted fan gas turbine engine generally 
indicated at 10 is of mainly conventional configuration. It comprises a 
nacelle 11 which encloses a core unit 12 and a propulsive fan 13 driven by 
the core unit 12. The core unit 12 includes intermediate and high pressure 
compressors 14 and 15 respectively, combustion equipment 16, and high, 
intermediate and low pressure turbines 17, 18 and 19 respectively. 
Three hollow concentric shafts 20, 22 and 23 interconnect the various 
turbines 17, 18 and 19 with the compressors 14 and 15 and the fan 13. More 
specifically, the radially innermost shaft 20 interconnects the low 
pressure turbine 19 with the fan 13, the radially outermost shaft 22 
interconnects the high pressure turbine 17 with the high pressure 
compressor 15 and the mid-shaft 23 interconnects the intermediate pressure 
turbine 18 with the intermediate pressure compressor 14. 
Bearing support structures 21, 24, 25 and 26, each including a series of 
radially extending struts, are positioned at the front, mid-region and 
rear of the core unit 12 to support the shafts 20, 22 and 23 on 
conventional roller bearings. Thus the mid-region bearing support 
structures 21 and 25 provide support for the radially outermost and 
mid-shafts 22 and 23 while the front and rear bearing support structures 
24 and 26 provide support for the mid-shaft 23 and the radially innermost 
shaft 20. 
The engine 10 operates in the conventional manner so that air entering the 
nacelle 11 at its upstream end (the left hand end when viewed in FIG. 1) 
is partially compressed by the fan 13. The air exhausted from the fan 13 
is divided by the upstream edge 27 of the core unit 12 into two concentric 
flows. The outermost flow of air is directed into an annular duct 28 
defined between the inner surface of the nacelle 11 and the outer surface 
of the core unit 12. That airflow is then exhausted to atmosphere at the 
downstream end of the nacelle 11 to provide propulsive thrust. 
The innermost flow of air is directed into the core unit 12 where it is 
compressed further by the intermediate and high pressure compressors 14 
and 15. The air is then directed into the combustion equipment 16 where it 
is mixed with fuel and the mixture combusted. The resultant hot combustion 
products then expand through, and thereby drive, the high, intermediate 
and low pressure turbines 17, 18 and 19 before exhausting to atmosphere 
from the downstream end of the core unit 12 to provide additional 
propulsive thrust. The turbines 17, 18 and 19 drive the compressors 15 and 
14 and the fan 13 respectively. 
Although the engine 10 has been described as having three concentric drive 
shafts 20, 22 and 23, it will be appreciated that it could, if so desired, 
have only two shafts so that low and intermediate pressure compressor and 
turbine systems are mounted on a common shaft. 
Conventionally, the nacelle 11 is provided within its walls with a gearbox 
which is driven by the radially outermost shaft 22 via a radially 
extending auxiliary drive shaft extending through one of the mid-region 
bearing support structures 21 and 25 and across the fan duct 28. However, 
in the case of the present invention, such a gearbox is omitted. Instead, 
the wall of the core unit 12 has positioned within it two similar 
electrical motor/generator units 29 and 30. 
The first motor/generator 29 is positioned in the upper region of the core 
unit 12. It is driven by the engine mid-shaft 23 via a drive shaft 31 that 
extends through one of the radially extending struts in the bearing 
support structure 25. Electrical power generated by the motor/generator 29 
during normal engine operation is used primarily to power the electrical 
systems that are specific to the engine 10 such as its full authority 
digital electric control system 32 located in the wall of the nacelle 11. 
In addition, electrical power from the motor/generator 29 is used to power 
engine accessories that are not normally electrically powered such as the 
variable inlet guide vanes, the compressor bleed valves and the thrust 
reverser 33. Conventionally thrust reversers are hydraulically powered in 
the conventional manner using hydraulic pumps driven by the main engine 
gearbox. 
Any excess power generated by the motor/generator 29 is directed, as 
required, to the main power supply system for the aircraft on which the 
engine 10 is operationally mounted. 
The second motor/generator 30 is positioned in the lower region of the core 
unit 12. It is driven by the radially outermost engine shaft 22 via a 
driveshaft 34 located in another of the struts in the bearing support 
structure 25. Electrical power generated by the second motor/generator 30 
during normal engine operation is primarily utilised as standby power for 
the aircraft on which the engine 10 is mounted. However, the primary 
function of the second motor/generator 30 is as a means for starting the 
engine 10. Thus when it is desired to start the engine 10, electrical 
power from a suitable external source is directed to the second 
motor/generator 30. This causes rotation of the outermost shaft 22 
interconnecting the high pressure compressor 15 and turbine 17. Sufficient 
air is compressed by the high pressure compressor 15 to sustain the 
combustion process within the combustion equipment 16 and thereby bring 
about the commencement of engine operation. 
The motor/generator 30 additionally drives a small gearbox 35 which if 
required provides power for pumping fuel and oil within the engine 10. 
The primary source of electrical power for the aircraft upon which the 
engine 10 is mounted is an electrical motor/generator 36 positioned within 
the downstream bearing support structure 26 at the downstream end of the 
innermost engine shaft 20. Powering the primary aircraft electrical 
motor/generator 36 by the innermost engine shaft 20 has the advantage of 
providing an emergency source of electrical power in the event of a 
failure of the engine 10. If the engine 10 ceases operation during 
aircraft flight, air flowing over the fan 13 causes rotation of the fan 13 
by the "windmill" effect, and hence operation of the electrical generator 
36. The advantage of this is that the aircraft no longer requires the 
provision of a separate ram air turbine to cope with the loss of 
electrical power resulting from engine failure. Such ram air turbines 
impose undesirable weight penalties on the aircraft. 
The gas turbine engine power system described with reference to FIG. 1 
provides useful advantages over a conventional ducted fan gas turbine 
engine. Elimination of the main engine gearbox and two electrical 
generators driven by the gearbox means that the shape of the nacelle 11 
can be optimised without the constraints imposed by the presence of the 
gearbox. Multiple sources of electrical power generation within the engine 
ensure that adequate electrical power generation capacity is available in 
the event of the failure of one of the generators. A further benefit is 
that power can be transferred between the main engine shafts 20, 22 and 23 
in the manner described in our co-pending PCT patent application no. WO 
95/02120. 
Further benefits may be enjoyed if the ducted fan gas turbine engine is of 
the type which is shown in FIG. 2. The basic components of the engine 37 
shown in FIG. 2 are the same as those of the engine shown in FIG. 1. 
Consequently components common to both engines 10 and 37 are given the 
same identification numbers and will not therefore be described in detail. 
The major difference between the ducted fan gas turbine engine 37 shown in 
FIG. 2 and that 10 shown in FIG. 1 is in the bearing arrangements for 
supporting the main engine shafts 20, 22 and 23. In the case of the engine 
37, all of the bearings supporting the main engine bearing shafts 20, 22 
and 23 are of the electromagnetic type. 
More specifically, the front bearing support structure 24 carries two 
spaced apart electromagnetic bearings 38 and 39. The upstream 
electromagnetic bearing 38 supports the upstream end of the radially 
innermost shaft 20. The downstream end of the radially innermost shaft 20 
is supported by a further electromagnetic bearing 40 which is carried by 
the downstream bearing support structure 26. The upstream electromagnetic 
bearing 38 is so configured that it additionally functions as an 
electrical generator and an electric motor. Electricity generated by the 
bearing 38 constitutes the primary source of electricity for the aircraft 
upon which the engine 37 is mounted. 
The other electromagnetic bearing 39 carried by the front bearing support 
structure 24 supports the upstream end of the mid-shaft 22. The 
electromagnetic bearing 39 is also of the type which is so configured that 
it also functions as both an electrical generator and an electric motor. 
Electricity generated by the bearing 39 is used in the main for providing 
all of the electrical supply requirements of the engine 37. The downstream 
end of the mid-shaft 22 is supported by an electromagnetic bearing 41 
carried by the bearing support structure 21. 
The mid-region bearing support structure 21 also carries an electromagnetic 
bearing 42 which supports the downstream end of the radially outer shaft 
22. The upstream end of the radially outer shaft 22 is supported by an 
electromagnetic bearing 43 carried by the other mid-region bearing support 
structure 25. Like the electromagnetic bearings 38 and 39, the 
electromagnetic bearing 43 is of the type which also functions both as an 
electrical generator and an electric motor. Electricity generated by the 
bearing 43 is utilised as a standby source of electrical power for the 
aircraft upon which the engine 37 is mounted. The bearing 43 functions as 
an electric motor during engine starting when it is used to rotate the 
radially outer shaft 22 in the manner described earlier. 
As in the case of the engine 10 described earlier, the provision of 
electrical generators and electric motors on all of the shafts 20, 22 and 
23 ensures that power transfer between the shafts 20, 22 and 23 is 
possible in the manner described in our co-pending PCT application Ser. 
No. WO95/02120. 
The use of electromagnetic bearings in the engine 37 brings important 
benefits in terms of overall engine operating efficiency. Thus it permits 
the total elimination of a conventional oil lubrication system with its 
problems of zoning and screening. It also permits the introduction of such 
desirable features as programmed control of shaft end loads, shaft bow 
effects following engine shut-down and subsequent rapid restart, shaft 
dynamics and imbalance effects and compressor and/or turbine blade tip 
clearances through axial shaft location. 
If electromagnetic bearings are utilised in a gas turbine engine as is the 
case with the engine 37, it becomes possible to integrate the operation of 
the engine with that of the auxiliary power unit carried by the aircraft 
which incorporates the engine 37. The manner in which this can be achieved 
can be seen more easily if reference is now made to FIG. 3. In FIG. 3 the 
engine 37 is shown in simplistic schematic form as is the auxiliary power 
unit 44. The auxiliary power unit 44 is of the conventional aircraft 
mounted gas turbine type. 
Before the engine 37 can be started, it is, of course, necessary to direct 
an electric current to the electromagnetic bearings 38, 39, 40, 41, 42 and 
43 in order to levitate them. The electric current may be derived from a 
ground source 45, a battery carried by the aircraft or from a conventional 
electric generator associated with the auxiliary power unit 44. 
Irrespective of its source, the electrical supply is directed along a line 
45 to each of the electromagnetic bearings 38, 29, 40, 41, 42 and 43. As 
soon as bearing levitation has been achieved, a further ground electrical 
supply 46 is directed through a line 47 to the electromagnetic bearing 43 
supporting the radially outer shaft 22. This causes the electromagnetic 
bearing 43 to function as an electric motor and thereby bring about 
rotation of the radially outer shaft 22. 
When the radially outer shaft 22 reaches a rotational speed sufficient to 
sustain operation of the engine 37, fuel is supplied to the combustion 
equipment 16 and ignition is commenced. As soon as effective engine 
operation is achieved, the electrical generators incorporated into the 
electromagnetic bearings 38, 39 and 43 provide sufficient electrical power 
to sustain the electrical demands of the engine 37 and of aircraft upon 
which it is mounted. Consequently the ground supplies of electricity 45 
and 46 or aircraft battery as the case may be can at this point be 
discontinued. 
During normal operation of the engine 37, the electrical generator 
incorporated into the electromagnetic bearing 38 supporting the radially 
inner shaft 20 provides the main electrical supply for the aircraft 
carrying the engine 37 through the line 48. The backup electrical supply 
for the aircraft is provided through the line 47 by the electrical 
generator incorporated into the electromagnetic bearing 43 supporting the 
radially outer shaft 22. Finally the electrical generator incorporated 
into the electromagnetic bearing 39 supporting the mid-shaft 23 provides 
the electrical supply for all of the electrical items associated with the 
engine 37, including, of course, its electromagnetic bearings. 
The auxiliary power unit 44 is arranged to operate at least at all times 
that the main engine 37 is operating. This is in complete contrast to 
normal auxiliary power unit operation in which the unit ceases operation 
as soon as the main engine is operational. It provides two outputs 49 and 
50. The output is of compressed air and is directed to the aircraft in 
order to provide a pressurised air supply for the aircraft cabin. 
Consequently, it is not necessary to tap air from the compressor system of 
the engine 37 to provide cabin air for the aircraft. The compromise in 
engine operating efficiency resulting from tapping air from the engine 
compressor system is thereby avoided. 
The second output 50 from the auxiliary power unit 44 is of pressurised 
hydraulic fluid. This is to power hydraulically driven accessories on the 
aircraft if such accessories are carried by the aircraft. 
In the event of a failure of the engine 37, the auxiliary power unit 44 
can, if desired, provide a continuous supply of electrical power to the 
electromagnetic bearings 38-43, thereby permitting the engine 37 to 
continue to rotate and generate electricity by the "windmill" effect 
described earlier. Alternatively, in a multi-engined aircraft electrical 
power generated by the other engine or engines could be used to levitate 
the electromagnetic bearings 38-43. 
It will be seen therefore that the integration of the gas turbine engine 37 
with the auxiliary power unit 44 provides a gas turbine engine system that 
brings about great flexibility in the way in which it is used to power the 
aircraft on which it is mounted and the various accessories with which it 
is associated. Thus, together, the turbine engine 37 and the auxiliary 
power unit 44 provide a number of different power sources that can be 
re-distributed between the engine and aircraft in accordance with the most 
efficient manner of operating the engine/aircraft combination. This means 
that the potential for overall engine propulsion efficiency improvements 
is high in view of the potential for using more aggressive engine cycle 
choices. Additionally, there is further potential for engine cycle 
performance and control system improvements through the use of 
electromagnetic bearings and the transfer of power between the various 
shafts 20, 22 and 23 of the engine 37. Weight and cost savings are further 
benefits to be enjoyed through systems integration. 
Although the engine 37 described above is of the type having three main 
shafts 20, 22 and 23, the present invention is applicable in its broadest 
aspect to two shaft engines of the type 51 shown in FIG. 4. The general 
features of the gas turbine engine/auxiliary power unit arrangement shown 
in FIG. 4 are very similar to those already described with reference to 
FIG. 3 and so common parts are depicted by common reference numerals. The 
only difference of significance is that the absence of a mid-shaft 23 
means that electrical power for the engine 51 is derived from the radially 
outer shaft 22 through the line 47. Line 47 still performs its other 
function of providing back-up electrical power for the aircraft.