Power plants – Reaction motor – Air passage bypasses combustion chamber
Reexamination Certificate
1999-07-14
2001-07-17
Thorpe, Timothy S. (Department: 3746)
Power plants
Reaction motor
Air passage bypasses combustion chamber
Reexamination Certificate
active
06260352
ABSTRACT:
BACKGROUND AND SUMMARY OF THE INVENTION
The present invention relates to a turbofan aircraft engine having a core engine, which is arranged in a gondola, and a fan which delivers a cold air current through a bypass flow duct to a mixer within which a low-pressure turbine cone is provided and on which the hot-gas current emerging from the low-pressure turbine of the core engine is mixed with the cold-air current. A portion of the cold-air current is branched off by an ejector effect and arrives, as a cooling air current for the low pressure turbine housing, in a ring gap between the housing and a covering surrounding this housing. A portion of this cooling air current is guided by hollow outlet guide blades of a follower guide wheel of the low-pressure turbine for sealing off the rearmost ring gap between the rotor and the stator of the low-pressure turbine, while the largest portion of the cooling air current reaches the ejector.
Reference is made to German Patent Document DE 33 24 347 C2 as well as to the applicant's BR 710 engine for examples of the technical environment.
A sufficient cooling of the low-pressure turbine housing can basically be achieved by certain characteristics. The state of the art known from the BR 710 aircraft engine is illustrated in FIG.
3
and will now be explained briefly. With respect to the overall arrangement in the aircraft engine, reference is partially made in this case also to
FIG. 1
which otherwise, however, shows the design according to the invention.
FIG. 1
is a partial longitudinal sectional view of a turbofan aircraft engine.
In
FIG. 1
, reference number
1
indicates the gondola of a turbofan aircraft engine, inside of which the so-called core engine
2
is concentrically arranged. In the downstream end area of the core engine
2
, as is customary, the low-pressure turbine
3
is situated and is adjoined by a low-pressure turbine cone
4
, while the fan
5
is provided upstream of the core engine
2
. This fan
5
is driven by the turbine shaft of the aircraft engine, which is not shown in detail, and delivers a cold-air current
6
through a bypass flow duct
7
situated between the gondola
1
and the core engine
2
.
Between the low-pressure turbine cone
4
and the interior wall of the gondola
1
, a ring-shaped mixer
8
is concentrically provided. This ring-shaped mixer may be constructed, for example, as a blossom-type mixer. The hot-gas current
9
coming out of the low-pressure turbine
3
is mixed on or in this mixer with the cold-air current
6
.
The low-pressure turbine housing is marked with reference number
10
in FIG.
1
. This low-pressure housing
10
is surrounded by a covering
11
of the low-pressure turbine
3
, and a ring gap
12
is situated between the covering
11
and the low-pressure turbine housing
10
.
A cooling air current
13
is guided through this ring gap
12
. This cooling air current is branched off the cold-air current
6
and is used particularly for cooling the low-pressure turbine housing
10
.
As illustrated in the known construction shown in
FIG. 3
, this cooling air current
13
is delivered, that is, branched off the cold air current
6
by an ejector effect. In the known construction, the downstream end area
11
a
of the covering
11
is constructed in the manner of a jet for this purpose; that is, in this end area
11
a,
the covering
11
together with the mixer inlet
8
a
forms a jet ring gap acting as an ejector
14
. By means of this ejector
14
, the cooling air current
13
is therefore entrained by the cold air current
6
flowing between the interior wall
1
a
of the gondola and the mixer
8
.
In order to prevent a so-called hot-gas entry (this term describes the occurrence of hot gas from the turbine annulus reaching the turbine shaft from the blades), in the known art, a fraction of the cooling air current
13
is guided through the hollow outlet guide blades
15
of the low-pressure turbine follower guide wheel of the core engine
2
, as a whole marked with reference number
16
, for sealing off the rearmost ring gap
17
between the rotor
18
and the stator
19
of the low-pressure turbine
3
. The corresponding flow of a fraction of the cooling air current
13
is indicated by the arrow
20
.
Based on 100% cooling air, which enters the ring gap
12
as the cooling air current
13
, for example, 5% reaches the ring gap
17
in the direction indicated by the arrow
20
, while the remaining 95% of the cooling air is removed by the ejector
14
in the direction indicated by the arrow
21
in the known construction.
This solution for cooling the low-pressure turbine housing
10
is basically acceptable, but when the cooling air current according to arrow
21
is reintroduced into the bypass flow duct
7
, flow losses occur and the specific engine consumption can be increased. In addition, for certain applications, it may be desirable to achieve an even stronger cooling of components of the low-pressure turbine
3
.
It is therefore an object of the present invention to provide guiding of the cooling air current on a turbofan aircraft engine which is improved with respect to the known prior art.
This object is achieved by having the largest portion of the cooling air current arrive through the hollow outlet guide blades in the interior of the low-pressure turbine cone which has a jet-type outlet opening at its downstream free end section so that this downstream end section acts as an ejector by which the cooling air current is admixed to the air current coming from the mixer. Advantageous further developments are also claimed.
REFERENCES:
patent: 2789416 (1957-04-01), Mirza
patent: 4175640 (1979-11-01), Birch et al.
patent: 4214441 (1980-07-01), Mouritsen et al.
patent: 5184459 (1993-02-01), McAndrews
patent: 5216879 (1993-06-01), Zysmaan
patent: 3210199 (1982-11-01), None
patent: 3940423 (1990-06-01), None
patent: 3324347 (1993-04-01), None
patent: 19524731 (1997-01-01), None
patent: 0514119 (1992-11-01), None
patent: 623615 (1949-05-01), None
patent: 695482 (1953-08-01), None
patent: 814380 (1959-06-01), None
patent: 2149022 (1985-06-01), None
Haarmeyer Jörg
Negulescu Dimitrie
Crowell & Moring LLP
Rodriguez W.
Rolls-Royce Deutschland Ltd & Co KG
Thorpe Timothy S.
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