Aircraft engine with pressure exchanger

In an aircraft engine with a fan or propeller drive, in which at least two shafts (1, 2) disposed to rotate coaxially and independently of one another respectively support a separate compressor (3, 4) and a separate turbine (5, 6), wherein a suitable device for combusting the gases is disposed between the compressor (4) disposed last in the flow direction and the first turbine (6), this device for combusting the gases comprises a dynamic pressure machine (9) known per se that operates with isochoric combustion and has a cellular wheel that rotates between respectively a side part on the air side and a side part on the gas side, both provided with input and output openings, and has a number of cells in which a continuously repeating ignition and combustion process takes place. In addition, a high-pressure turbine (10) is disposed behind the dynamic pressure machine (9) on the same shaft (2) as the following medium-pressure turbine (6).

FIELD OF THE INVENTION 
The invention relates to an aircraft engine with a fan or propeller drive, 
in which at least two shafts disposed to rotate coaxially and 
independently of one another respectively support a separate compressor 
and a separate turbine, and a suitable device for combusting the gases is 
disposed between the compressor disposed last in the flow direction and 
the first turbine. The invention also relates to a method for operating 
this aircraft engine. 
BACKGROUND OF THE INVENTION 
Aircraft engines of this type are known. Aircraft engines with a dual-shaft 
design have two shafts seated to rotate coaxially and independently of one 
another, and each has respectively a separate compressor and a separate 
turbine. The combustion chamber is disposed between the medium-pressure 
compressor and the medium-pressure turbine, and in turbojet engines, the 
propulsion jet is disposed downstream of the medium-pressure turbine. In 
contrast, in aircraft engines with a propeller the propulsive power is 
essentially generated by means of the transmission of a shaft output to a 
propeller operating in atmospheric air. 
The embodiment of the by-pass flow of dual-flow turbojet engines typically 
includes the turbofan operating in the by-pass flow and whose blades can 
either form an extension of the compressor blades (front fan) or the 
turbine blades (aft fan) of the main rotor, or is driven by a separate 
rotor. The air flow entering into the engine is guided in a hot primary 
flow and in a second, cold secondary flow that is essentially independent 
of the primary flow, and the primary flow is surrounded coaxially and at 
least partly along its length by the secondary flow in annular or 
sheath-type fashion. In a bypass engine with a dual-shaft embodiment, this 
secondary flow is branched off from the primary air flow in a controlled 
manner and mixed again with the actual propulsion stream in the nozzle. 
The primary flow is compressed in the low-pressure compressor and 
subsequently in the medium-pressure compressor. Afterward, heat energy is 
supplied to it in the combustion chamber by means of the combustion of 
continuously injected fuel with the oxygen of the compressed air. Only as 
much energy is drawn from the compressed and heated air or combustion gas 
by the downstream turbines as is required to drive the compressors and the 
turbofan or the propeller. 
In comparison to conventional turbojet engines, dual-flow turbojet engines 
of this type have an increased propulsion effect, an improved ability to 
be controlled and a reduced specific fuel consumption, particularly at 
average flying speeds. Thus, any further improvement in the performance of 
aircraft engines of this type no longer appears to be possible. 
Stationary gas turbine systems are known in which a dynamic pressure 
machine operating with isochoric combustion is disposed between the 
compressor and the gas turbine. This dynamic pressure machine has a 
cellular wheel that rotates between respectively a side part on the air 
side and a side part on the gas side, both provided with input and output 
openings, and has a number of cells in which a continuously repeating 
ignition and combustion process takes place. If, during operation, the 
dynamic pressure machine produces a homogenous fuel distribution in the 
longitudinal direction of the cells (European Patent Disclosure EP 
0468083), problems occur when a low-pressure turbine and a high-pressure 
turbine are connected to this dynamic pressure machine, because both 
turbines are to be operated at the same optimum temperature if possible, 
although the gases are supplied to the high- and low-pressure turbines at 
different pressures. Therefore a change has been made to generate two 
partial quantities of gas enriched to different extents with fuel from 
compressed mass flows that are released from the cells after ignition and 
compression, wherein the partial quantity enriched to a lesser extent with 
fuel is supplied to the high-pressure turbine, and the second partial 
quantity enriched to a greater extent with fuel is supplied to the 
low-pressure turbine. It has been achieved by means of the different 
degrees of enrichment of the partial quantities with fuel, that neither 
overly hot gases are supplied to the high-pressure turbine, nor are overly 
cold gases supplied to the low-pressure turbine. This results in an 
improvement in the efficiency of the gas turbine systems. 
OBJECT AND SUMMARY OF THE INVENTION 
The object of the invention is to use suitable means and develop a method 
with which it is possible to increase the efficiency and the specific 
performance of aircraft engines of fan or propeller drive and in which at 
least two shafts seated to rotate coaxially and independently of one 
another respectively support a separate compressor and a separate turbine, 
and wherein a device for combusting the gases is disposed between the last 
compressor in the flow direction and the first turbine. 
This object is attained in accordance with the invention in that a dynamic 
pressure machine that operates with isochoric combustion is disposed 
between the last compressor in the flow direction and the first gas 
turbine, which has a cellular wheel that rotates between respectively a 
side part on the air side and a side part on the gas side, both provided 
with input and output openings, and has a number of cells in which a 
continuously repeating ignition and combustion process takes place, and 
that a high-pressure turbine is additionally disposed behind the dynamic 
pressure machine on the same shaft as the following medium-pressure 
turbine. With the method for operating the aircraft engine, the pressure 
conditions in either the compressors or in the turbines are adapted. 
It is useful when the dynamic pressure machine is disposed in such a way 
that its rotor free-wheels at a low rpm. 
It is also advantageous when the dynamic pressure machine can be driven by 
a planetary gear. 
The advantages of the invention can be seen in, among other things, 
low-pollutant combustion, an improved efficiency and an improvement in 
performance as compared with conventional aircraft engines with the same 
core engine. 
Exemplary embodiments of the invention are shown by way of a dual-shaft 
aircraft engine with axial flow-through.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
Only the elements essential for understanding the invention are shown. The 
flow direction of the working means is indicated by arrows. 
FIG. 1 shows a conventional aircraft engine of the dual-shaft type with a 
fan. The low-pressure compressor 3 with a fan 7 and the low-pressure 
turbine 5 are located on a common shaft 1. The medium-pressure compressor 
4 and the medium-pressure turbine 6 are seated on a common shaft 2. Shafts 
1 and 2 are seated to rotate coaxially and independently of one another. 
The combustion chamber 11 is disposed between the medium-pressure 
compressor 4 and the medium-pressure turbine 6. The propulsion jet 8 is 
connected to the low-pressure turbine 5. 
The air flow entering into the aircraft engine enters a turbofan 7, is 
compressed slightly, for instance from approximately 1 to approximately 2 
bar, and split into a primary air flow with a mass flow m.sub.1 and an 
independent, secondary air current with a mass flow m.sub.2. The primary 
air flow m.sub.1 is surrounded coaxially and along its entire length in a 
sheath-like manner by the secondary flow. The primary air flow is 
compressed in the compressors 3 and 4 to approximately 6 and approximately 
30 bar and conducted into the combustion chamber 11. Heat energy is 
supplied there by means of the combustion of continuously injected fuel 
with the oxygen of the compressed air. The medium-pressure turbine 6 
(entry pressure approx. 30 bar) disposed in front of the propulsion jet 8 
and the low-pressure turbine 5 (entry pressure approx. 10 bar) draw only 
enough energy out of the air as is needed to drive the medium-pressure 
compressor 4, the low-pressure compressor 3 and the turbofan 7. The 
remaining supplied heat energy is utilized for propulsion as kinetic 
energy of the propulsion stream emitted from the propulsion jet 8. 
An exemplary embodiment of the invention in which the pressure conditions 
are adapted in the compressors is shown in FIG. 2. It differs from the 
conventional aircraft engine in FIG. 1 in that a dynamic pressure machine 
9 with isochoric combustion is disposed in place of the combustion chamber 
11 (FIG. 1 ). The rotor 20 of the dynamic pressure machine free-wheels, 
but can also be driven with a planetary gear 22, as illustrated 
schematically in FIG. 5. For freewheeling, the gear system 22 is omitted, 
and the bearings 24 alone support the rotor 20 of the dynamic pressure 
machine. In comparison with FIG. 1, in FIG. 2 a high-pressure turbine 10 
is additionally disposed in front of and on the same shaft 2 as the 
medium-pressure turbine 6. As in the conventional engine in accordance 
with FIG. 1, the air flow is divided into a primary air flow with a mass 
flow m.sub.1 and a secondary air flow with a mass flow m.sub.3. The 
secondary air flow m.sub.3 bypasses the compressors 3, 4, the dynamic 
pressure machine 9 and the turbines 10, 6 and 5. After flowing through the 
turbofan 7, the mass flow m.sub.1 is compressed in the low-pressure 
compressor 3 from approximately 2 to approximately 4 bar, for example, 
then subsequently compressed from approximately 4 to approximately 30 bar 
in the medium-pressure compressor, and at this pressure reaches the intake 
opening of the dynamic pressure machine 9. In the dynamic pressure machine 
9, the compressed air is enriched with fuel in such a way that three 
different layers form: the first layer contains no fuel, the second layer 
contains a small quantity and the third layer contains a large quantity of 
fuel. The first layer is used as cooling air to cool the walls of the 
dynamic pressure machine 9 and the high-pressure turbine 10. After 
combustion of the fuel-air mixture in the cells of the cellular wheel of 
the dynamic pressure machine 9, The second layer, which was comprised of a 
lean fuel-air mixture, is supplied to the high-pressure turbine 10 from 
the dynamic pressure machine 9 through a first outlet opening, while the 
third layer, which was comprised of the richer fuel-air mixture before 
combustion, is supplied to the medium-pressure turbine 6 through a second 
outlet opening. The exact mode of operation of the dynamic pressure 
machine with isochoric combustion has already been described in 
Applicant's previous patent applications, so details are omitted here. 
Through the use of the dynamic pressure machine 9, a portion of the gas is 
compressed further so that the second layer enters the high-pressure 
turbine 10 at a pressure of approximately 60 bar, the turbine being 
additionally disposed, in contrast to the prior art. The third layer is 
supplied to the medium-pressure turbine 6 at a pressure of approximately 
30 bar. The gas expands in the turbines 10 and 6. Thanks to the more 
extensive enrichment of the third layer with fuel, this layer has the same 
temperature as the second layer, although the third layer expands more. By 
means of this, both the high-pressure turbine 10 and the medium-pressure 
turbine 6 can be operated at the same optimum temperature. The gas 
subsequently enters the low-pressure turbine 5 at a pressure of 
approximately 10 bar, and last of all the propulsion jet 8. 
In comparison to the prior art, the use of the dynamic pressure machine 9 
and the high-pressure turbine 10 results in an improved efficiency and 
performance in a constant core engine. The bypass ratio m.sub.3 /m.sub.1 
of the engine in accordance with the invention is greater than the bypass 
ratio m.sub.2 /m.sub.1 of a comparable conventional aircraft engine, 
because a larger fan can be driven by the additional capacity, and a 
larger secondary mass flow m.sub.3 can be realized. 
The latter also applies to the exemplary embodiment shown in FIG. 3, in 
which the pressure conditions are adapted in the turbines. The compressors 
3, 4 are designed in the same manner as in the conventional aircraft 
engine in FIG. 1. The dynamic pressure machine 9 with isochoric combustion 
is disposed in place of the combustion chamber 11 and followed by an 
additional high-pressure turbine 10. The mode of operation of the dynamic 
pressure machine 9 is the same as described above in FIG. 2. Unlike in the 
example in FIG. 2, the gas is not greatly expanded in the medium-pressure 
turbine 6, for example from approximately 30 to approximately 14 bar, but 
instead to approximately 10 bar. The low-pressure turbine 5 connected to 
the medium-pressure turbine 6 therefore has an additional capacity that is 
utilized to drive a larger turbofan 7. Thus the mass flow m.sub.3 is also 
increased here, and a greater propulsion power can be achieved with this 
aircraft engine. 
In the exemplary embodiment shown in FIG. 4, the additional capacity 
achieved by means of the installation of the dynamic pressure machine 9 
and the additional high-pressure turbine 10 is used to increase the 
transmission of the shaft output to the propeller 12, thus improving 
propulsion. 
To summarize, the advantages of the invention lie in the fact that, in 
comparison to conventional aircraft engines, with a constant core engine 
either an improved efficiency and better performance are achieved, or the 
same performance is achieved with smaller engines.