Variable cycle propulsion unit for aircraft

A variable cycle propulsion unit comprises at least two turbojet engines of the bypass type in which the low pressure compressor of each engine is in two parts connected by a shaft, and the flow of compressed air issuing from the first part of the low pressure compressor of a first of the engines is supplied to the second part of the low pressure compressor of a second of the engines and the flow of compressed air issuing from the first part of the low pressure compressor of the second engine is supplied to the second part of the low pressure compressor of the first engine. At full throttle operation the combustion chambers of the two engines are both ignited, and at operation at reduced engine speeds the combustion chamber of one of the engines is extinguished and the engine windmills while the other engine continues to run normally.

FIELD OF THE INVENTION 
The present invention relates to a variable cycle propulsion unit, 
particularly for a military fighter aircraft, as well as to an 
advantageous method of operating the unit. 
BACKGROUND OF THE INVENTION 
One development in modern fighter aircraft is a tendency to increase the 
installed power. Indeed, manoeuvrability requirements necessitate the 
provision of highly powered aircraft. The search to reduce the infra-red 
radiation of aircraft without reducing performance leads to an increase in 
installed power without after-burn. This is also a consequence of the 
search for versatility, such as in STOL/VTOL applications. 
The outcome, for a given technological development level, is an increase of 
engine mass which imposes a reduction of overall fuel capacity of up to 20 
to 25% in order to keep constant the total mass of the aircraft on 
take-off. 
In the known engines used nowadays, of the bypass turbojet type, 
suppressing the after-burn makes it possible to obtain, when operating at 
full throttle, a substantial reduction in fuel consumption which may 
exceed 100%, but, at constant cruising thrust, the specific fuel 
consumption increases with the size of engine. 
A representation is given in FIG. 1 of the attached drawings, of a curve 
showing the variation of the ratio Cs/Csm (as ordinate) of the specific 
consumption at partial throttle and full throttle as a function of the 
ratio F/Fm (as abscissa) of the corresponding thrusts at partial throttle 
and maximum throttle. On this curve, point A corresponds to an engine 
equipped with after-burn, and point B to an engine without after-burn 
supplying the same thrust on take-off, and therefore of greater size. It 
will be observed that for an increase of engine size enabling an increase 
of thrust .DELTA.F/F of 50% to be obtained, the corresponding increase of 
specific consumption when cruising .DELTA.Cs/Cs is 26.5%. 
Indeed, from the analyses carried out the following requirements appear 
difficult to reconcile, namely: 
to obtain a powerful thrust at full throttle, it is necessary to operate 
with a low bypass ratio and to use an engine with a large high pressure 
body; 
whereas, to obtain a low fuel consumption when cruising, it is necessary to 
operate with a high bypass ratio and to use an engine with a small high 
pressure body. 
SUMMARY OF THE INVENTION 
It is thus an object of the invention to provide a propulsion unit which 
meets these requirements without suffering from the drawbacks of the 
currently known variable cycle engines. 
To this end, according to the invention there is provided a variable cycle 
propulsion unit of the turbojet type, particularly for a military fighter 
aircraft, comprising at least two base engines of the bypass type, each of 
said base engines including, from upstream to downstream relative to the 
normal direction of flow of gases in the engine, a low pressure 
compressor, a high pressure compressor, a combustion chamber, at least one 
turbine driving said compressors, and an ejection nozzle assembly, wherein 
said low pressure compressor of each of said base engines comprises a 
first part including a first stage of rotor blades, a second part 
including at least one second stage of rotor blades, an inlet stage of 
fixed guide vanes preceding said at least one second stage of rotor 
blades, and an outlet stage of fixed guide vanes following said at least 
one second stage of rotor blades, and a shaft interconnecting said first 
and second stages of rotor blades, and wherein said propulsion unit 
includes first means for delivering the compressed air provided by said 
first part of said low pressure compressor of a first of said base engines 
to said second part of said low pressure compressor of a second of said 
base engines, and second means for delivering the compressed air provided 
by said first part of said low pressure compressor of said second base 
engine to said second part of said low pressure compressor of said first 
base engine, thereby effecting a permanent cross-over of the flows between 
said first and second base engines of said propulsion unit. 
Preferably, each of said first and second means comprises a first toric 
volute for receiving directly the compressed air delivered by said first 
part of said low pressure compressor of the respective one of said first 
and second base engines, a second toric volute preceding said second part 
of said low pressure compressor of the other of said first and second base 
engines, and a duct connecting said first toric volute to said second 
toric volute. By arranging for a simple pivotal movement of the assembly 
of volutes through 180.degree. about the engine axis, the two engines may 
be made identical and will accordingly be interchangeable. 
The invention also provides a method of operating the propulsion unit 
wherein, at full throttle, said combustion chambers of said first and 
second base engines are ignited and operate the respective high pressure 
bodies, corresponding to operation of a large high pressure body with a 
low bypass ratio, and, at reduced throttle, said first engine is stopped 
and its combustion chamber extinguished, while said second engine 
continues to operate normally with said first part of said low pressure 
compressor of said second engine delivering air to said first engine and 
causing said first engine to windmill, corresponding to operation of a 
small high pressure body with a high bypass ratio.

DESCRIPTION OF THE PREFERRED EMBODIMENT 
A propulsion unit in accordance with the invention and intended for a 
military fighter aircraft is formed by an even number of base engines 
associated in pairs. In the embodiment represented in FIGS. 2 to 4, the 
propulsion unit is formed of two associated base engines. Each engine is a 
turbojet engine of the bypass type. A first engine 1 thus includes, in a 
known and conventional manner, a high pressure primary unit comprising a 
high pressure compressor 2, a combustion chamber 3 and a high pressure 
turbine 4, and, in a similar manner, the second engine 1a comprises a high 
pressure compressor 2a, a combustion chamber 3a and a high pressure 
turbine 4a. This primary unit of each engine is surrounded by an annular 
secondary duct 5, 5a respectively, termed the "cold flow duct", and each 
engine comprises at its downstream end (upstream and downstream or front 
and back being defined relative to the normal direction of the flow of 
gases through the engine) an ejection nozzle assembly 6, 6a which is of 
known construction and which is not shown in detail in the drawings. 
The turbojet engines may be of the single spool type, or, as in the example 
shown in the drawings, of the twin spool type. In the latter case, as is 
diagrammatically shown in FIG. 2, the high pressure compressor 2 or 2a and 
the high pressure turbine 4 or 4a are connected by means of a shaft 7 or 
7a, in each engine. A low pressure spool is driven by a low pressure 
turbine 8 or 8a connected by means of a shaft 9 or 9a to a low pressure 
compressor 10 or 10a. As mentioned earlier, it is desirable for military 
aircraft to avoid an increase of infra-red radiation, and accordingly the 
propulsion unit of the invention is preferably not provided with an 
after-burner system. However, provided it is not too substantial, an 
after-burner of known construction may be included in the propulsion unit 
of the invention if desired. 
An essential feature of the invention resides in the construction of the 
low pressure compressors 10 and 10a of the base engines of the propulsion 
unit. Each low pressure compressor 10, 10a is made in two parts. A first 
part 11, 11a is composed of the front part of the rotary assembly and 
comprises a first stage 12, 12a of rotor blades, and the second part 13, 
13a comprises at least one second stage 14, 14a of rotor blades preceded 
by a stage 15, 15a of fixed vanes forming an inlet guide and followed by a 
stage 16, 16a of fixed vanes forming an exit guide. A part of the air flow 
thus compressed, symbolized by the arrow 17, is supplied to the secondary 
duct 5 before being ejected, and another part of the compressed air flow, 
symbolized by the arrow 18, is supplied to the high pressure compressor 2 
and the corresponding primary unit. 
The two parts 11, 11a and 13, 13a of each low pressure compressor are 
spaced apart axially, and the hubs of the first and second stages of rotor 
blades 12, 12a and 14, 14a are connected by a shaft 19, 19a. The 
interconnection of the two base engines 1 and 1a forming the propulsion 
unit is effected at the level of the low pressure compressors 10 and 10a 
by a permanent crossing-over of their respective flows. Thus, the air 
flow, symbolized by the arrow 20, compressed by the first part 11 of the 
low pressure compressor 10 of the first engine 1 is conducted to the inlet 
of the second part 13a of the low pressure compressor 10a of the second 
engine 1a and, in accordance with the symmetrical arrangement chosen for 
the two engines, the air flow, symbolized by the arrow 21, compressed by 
the first part 11a of the low pressure compressor 10a of the second engine 
1a is conducted to the inlet of the second part 13 of the low pressure 
compressor 10 of the first engine 1. 
As shown by FIGS. 3 and 4, in this embodiment the said air flows 20 and 21 
are guided by two ducts 22 and 23 respectively. The first parts 11, 11a of 
the low pressure compressors 10, 10a each deliver into a respective toric 
volute 24, 24a which collects the whole of the flow and channels it to an 
outlet opening, which, in the case of the first engine 1 (i.e. the right 
hand engine when looking at the propulsion unit from the rear), is an 
opening 25 situated on the left and at the top of the engine, and in the 
case of the left hand engine 1a is an opening 25a situated to the right 
and at the bottom of the engine. The second parts 13, 13a of the low 
pressure compressors are each served by a supply volute 26, 26a situated 
upstream of the inlet guide 15, 15a and having an intake opening 27, 27a. 
In the case of the first (or right hand) engine 1, this opening 27 is 
situated on the left and at the bottom of the engine, and is connected by 
means of the said duct 23 to the outlet opening 25a of the volute 24a of 
the second engine 1a. In the case of the second (or left hand) engine 1a, 
the intake opening 27a is situated on the right and at the top of the 
engine, and is connected by the said duct 22 to the outlet opening 25 of 
the first engine 1. The two ducts 22 and 23 are designed so that they can 
be removed, and are preferably made from a flexible material. The unit 
formed by the two volutes 24 and 26 of the right engine 1 is identical 
with that formed by the two volutes 24a and 26a of the left engine 1a, and 
the effective conversion of the right engine into a left engine, and vice 
versa may be obtained simply by pivoting the entire assembly of volutes 
through 180.degree. around the axis of the propulsion unit. 
The propulsion unit which has just been described with reference to FIG. 2 
to 4 makes it possible to achieve the aims of the invention and the 
advantages mentioned earlier, particularly on applying the principle of 
operation or the modus operandi in accordance with the invention which is 
described below. 
Two principal operating modes of the propulsion unit are used. Firstly, 
when operating at full throttle, the two associated engines 1 and 1a 
operate as conventional bypass turbojet engines. Each high pressure 
compressor 2, 2a is supplied by the second part of the corresponding low 
pressure compressor 10, 10a and supplies the combustion chamber 3, 3a 
which is ignited and despatches the burnt gases through the turbines 4, 4a 
and 8, 8a to the ejection nozzle 6, 6a. Recourse is had to this mode of 
operation each time a powerful thrust is required, such as on take-off or 
in combat. The propulsion unit can in this case be regarded as equivalent 
to a single engine having the characteristics of a low bypass ratio and a 
large high pressure body or spool. 
Secondly, when operating at reduced throttle, such as during escort flights 
or in cruising or waiting phases, one of the two engines, for example the 
first engine 1, is shut down and its combustion chamber 3 extinguished, 
while the second engine 1a continues to be operated normally as described 
above. In this case, the first part 11a of the low pressure compressor 10a 
of the second engine 1a continues to supply the second part 13 of the low 
pressure compressor 10 of the first engine 1, which thus delivers a cold 
flow causing the first engine to windmill. The first part 11 of the low 
pressure compressor 10 of the first engine 1 thus windmills and continues 
to supply the second part 13a of the low pressure compressor 10a of the 
second engine la as well as its primary body. 
In order to ensure the compatibility of the flow rates between the first 
parts 11 and 11a and the second parts 13 and 13a of the low pressure 
compressors 10 and 10a, a device 28, 28a and 29, 29a for adjusting the 
pitch of the fixed blades is provided in a known manner on each guide 
stage 15, 15a and 16, 16a respectively of the low pressure compressors 10 
and 10a. Alternatively, a discharge device of known construction may be 
used to regulate the flows of the low pressure compressors 10 and 10a. 
In this second operating mode the propulsion unit may be regarded as 
equivalent to a single engine having the characteristics of a high bypass 
ratio and a small high pressure body or spool, as its size is effectively 
reduced by half relative to the first operating mode. This method of 
operation permits, in particular, the achievement of a lower compression 
ratio at low pressure, a higher bypass ratio and a lower overall 
compression ratio. This, in turn, achieves one of the objectives sought, 
namely the lowering of the specific fuel consumption level. Depending on 
the type of engine, with or without after-burner, taken as a basis for 
comparison, the gains in specific consumption obtained may range from 25 
to 30% up to 40 or 45%. 
A further advantage of the propulsion unit which is the subject of the 
present invention is that it is not essential to couple the two engines to 
carry out development tests on the propulsion unit. A few slight 
modifications will suffice to achieve, from one engine only, a simulation 
of operation on full throttle, or a simulation of operation at reduced 
throttle. For this purpose, as shown diagrammatically in FIG. 5, operation 
at full throttle can be effected by mounting a connecting duct 30 between 
the outlet of the first part 11 and the inlet of the second part 13 of the 
low pressure compressor 10 of the engine selected for the tests. 
Alternatively, as shown diagrammatically in FIG. 6, to simulate operation 
at low throttle, a first adaptor 31 is mounted at the inlet of the second 
part 13 of the low pressure compressor 10 to provide the supply of air and 
simulate a loss of head equivalent to that resulting from the first 
compressor stage, and a second adaptator 32 is mounted at the outlet of 
the first part 11 of the low pressure compressor 10 to simulate the loss 
of head corresponding to flow in the second engine.