Note: Descriptions are shown in the official language in which they were submitted.
S~7~
BACKGROUND OF T~IE INVENTION
-
This invention relates to gas turbines and, more particularly,
to a concept for efficiently reducing the temperature of air used to cool high
temperature turbines in gas turbofan engines.
Modern aircraft gas turbofan engines operate at turbine inlet
air temperature levels which are beyond the structural temperature capabilities
of high temperature alloys. Hence, engine hot flow path components and, in
particular, turbine blades and vanes must be cooled in order to assure their
structural integrity in order to meet operating life requirements. It is well
understood that gas turbine engine shaft horsepower and specific fuel consump-
tion (which is the rate of fuel consumption per unit of power output) can be
improved by increasing turbine inlet temperature. In order to take advantage
of this potential performance improvement, modern turbine cooling technology
utilizes air-cooled, hollow turbine nozzle vanes and blades to permit operation
at inlet gas temperatures in excess of 2000 F (1094 C). In general, these
sophisticated methods of turbine cooling have utilized compressor discharge
or interstage bleed air as a coolant. However, the benefits obtained from
sophisticated air-cooling techniques are at least partially offset by the extrac-
tion of the necessary c~oling air from the propulsive cycle. It can be
appreciated that the cooling airflow rate required is a function of the hot gas
temperature, increasing with increasing hot gas temperature. Furthermore,
the compressor bleed air used for cooling must bypass the combustor and one
or more turbine stages, thus giving rise to a performance penalty proportion-
ate to the amount of cooling air utilized. More particularly, the air that is
bled from the compressor and used as cooling air for the turbine rotor blades
has had work done on it by the compressor, However, because it is normally
--1--
returned into the flow path gas stream downstream of the turbine nozzle, it
does not return its full measure of work to the cycle as it expands through the
. turbine. Additionally, the reintroduction of cooling air into the hot gas
stream produces a loss in gas stream total pressure. This is a result of the
momentum mixing losses associated with injecting relatively low total pressure
cooling air into a high total pressure gas stream. Thus, the greater the
amount of cooling air which is routed through the turbine blades, the greater
the losses associated with the coolant become on the propulsive cycle. Thus,
while turbine blade cooling has inherent advantages, it also has associated
therewith certain inherent disadvantages which are functions of the quantity of
cooling air used in cooling the turbine rotor blades.
It will, therefore, be appreciated that engine performance can
be increased by reducing the amount of cooling air required by the turbine.
Reducing the cooling airflow rate results in improved engine performance with
a consequent reduction in specific fuel consumption, the actual magnitude of
the cooling airflow rate and specific fuel consumption reductions which can be
realized being a function of the specific engine application.
One method of reducing the amount of cooling air required by
the turbine is to cool the cooling air entering the hot components. One widely
advocated method of cooling the cooling air is to utilize the heat sink capability
available in the engine fuel. In such a scheme, the relatively hot cooling air
is placed in heat exchange relationship with the relatively cool engine fuel,
thereby cooling the cooling air and heating the fuel. The energy extracted by
the fuel is reintroduced back into the propulsive cycle as the heated fuel is
burned in the combustor, thereby producing what has commonly been referred
to as a "regenerative engine. " While various studies indicate that fuel-air
-2 -
.
l~SZ~6
heat exchangers offer an advantage of small size and low weight, the fuels
currently used in aircraft engines (JP4, JP5) are limited in their heat sink
capacity, the available heat sink already being used largely to cool the engine
oil. To obtain an additional heat sink capacity to permit cooling of the coolingair would require the use of special fuels such as JP7 or JP9, which are
currently unavailable in commercial quantities. Additionally, the use of fuel ina fuel-air heat exchanger presents a potential fire hazard which may be unaccept-
able for commercial engine applications. It will, therefore, be appreciated
that another technique for cooling the cooling air is required in order to reduce
the coolant flow rate and thereby enhance overall engine performance.
SUMMARY OF THE INVENTION
. .
Accordingly, it is the primary object of the present invention to
provide for a reduction in the amount of cooling air required by the turbine of
a gas turbofan engine by reducing the temperature of the cooling air passing
therethrough in order to enhance overall engine performance.
This, and other objects and advantages, will be more clearly
understood from the following detailed descriptions, the drawing and specific
examples, all of which are intended to be typical of, rather than in any way
limiting on, the scope of the present invention.
Briefly stated, the above objects are obtained in an aircraft
gas turbofan engine by providing a heat exchanger wherein the turbine cooling
air and relatively cooler air from the fan bypass duct are maintained in heat
exchange relationship, thereby cooling the turbine cooling air. The turbine
cooling air is bled, for example, from the discharge of the compressor through
ports in the engine casing at various circumferential locations and is ducted tothe heat exchanger which is disposed inwardly of the fan bypass portion of the
--3--
S2~6
. .
gas turbine engine. The relatively cool fan bypass duct air is bled at the inner
wall of the fan duct into a diffuser where the dynamic head of the fan stream is
largely recovered. The fan bleed air is then ducted through the heat exchanger
into heat exchange relationship with the relatively warmer compressor
5 discharge bleed air, thereby absorbing heat from the cooling air, and returned
to the fan bypass duct, The cooled compressor discharge bleed air is then
routed to the high pressure turbine through the compressor rear frame struts
and is expanded through an expander nozzle prior to cooling the high pressure
turbine components. In an alternative embodiment of the present invention
10 where the space available for ducting the cooling air through the compressor
rear frame struts is limited, the cooling flow rates through the heat exchanger
may be reduced by increasing the magnitude of the cooling air temperature
reduction in the heat exchanger in direct proportion to the reduction in flow
rates. The resulting over-cooled cooling air i9 then mixed with uncooled
15 compressor discharge bleed air ahead of the expander nozzle to obtain the
cooling air temperature reduction necessary to cool the turbine.
Incorporation of this heat exchanger into an aircraft gas turbofan
engine permits a reduction in the quantity of air required for turbine cooling
and, thus, provides an improvement in engine performance. Conversely, an
20 increase in blade life can be achieved by maintaining the original coolant flow
rate but by reducing the temperature of the coolant, with essentially no further
degradation in engine performance.
BRIEF DESCRIPTION OF THE DRAWING
-
While the specification concludes with claims particularly
25 pointing out and distinctly clairning the subject matter which is regarded as
part of the present invention, it is believed that the invention wilI be more fully
--4--
~D .
11~7~
understood from the following description of the preferred embodiments which
are given by way of example with the accompanying drawing in which:
Figure 1 is a simplified cross-sectional view, in partial cutaway,
of an aircraft gas turbo~an incorporating the preferred embodiment of the
5 subject invention and illustrating the relationship of the heat exchanger to the
various other engine components;
Figure 2 is a simplified cross-sectional view of a portion of the
gas turbofan engine of Figure 1 depicting an alternative embodiment of the
cooling system of the present invention; and
Figure 3 graphically depicts the turbine relative cooling flow
rate and specific fuel consumption reductions as a function of the change in
cooling air temperature for the representative gas turbofan engine of Figure 1.
DESCRIPTION OF THE PREFERRED EMBODIMENT
Referring to the drawing wherein like numerals correspond to
like elements throughout, attention is first directed to Figure 1 wherein a
representative gas turbofan engine designated generally at 10, and which
incorporates the present invention, is diagrammatically shown. While it is
. recognized that turbofan engines are, by now, well known in the art, a brief
description of the operation of the engine will enhance appreciation of the
interrelationship of the various components in light of the invention soon to be
described. Basically, this engine may be considered as comprising a core
engine 12, a fan 14 including a rotatable stage of fan blades 16 (only one of
which is shown for clarity), and a fan turbine (not shown) downstream of the
core engine in the area generally depieted as 17 and which is interconnected to
the fan 14 by shaft 18, The core engine 12 includes an axial flow compressor
20 having a rotor 22. Air enters inlet 24 from the left of Figure 1 and is
--5--
ll~'S~7~
initially com~ressed by the fan blades 16. A first portion of this relatively
cool cornpressed air enters the fan bypass duct 26 defined, in part, by core
engine 12 and a circumscribing fan cowl or nacelle 28 and discharges through
a fan nozzle 30. A second portion of the compressed air enters core engine
5 inlet 32, is further compressed by the axial flow compressor 20 and is
discharged to a combustor 34 where it is mixed with fuel and burned to provide
high energy combustion gases which drive a core engine turbine 36. The
turbine 36, in turn, drives the rotor 22 through a shaft 38 in the usual manner
of a gas turbine engine. The hot gases of combustion then pass through and
10 drive the fan turbine which, in turn, drives the fan 14. A propulsive force is
thus obtained by the action of the fan 14 discharging air from the fan bypass
duct 26 through the fan nozzle 30 and by the discharge of combustion gases
from a core engine nozzle 40 defined, in part, by plug 42. The above descrip-
tion is typical of many present-day gas turbofan engines and is not meant to be
15 limiting to the present invention, as it will become readily apparent from the
following description that the present invention is capable of application to any
gas turbofan engine of the bypass variety and is not necessarily restricted to
use with the particular configuration depicted herein. The foregoing descrip-
tion of the operation of the engine depicted in Figure 1 is, therefore, merely
20 meant to be illustrative of one type of application,
It is also well understood that gas turbine engine shaft horse-
power and specific fuel consumption (which is the rate of fuel consumption per
unit of power output) can be improved by increasing the temperature at the
inlet tothe core engine turbine 36 (sometimes referred to ~s the "high pressure
25 turbine"). However, since modern aircraft turbofan engines operate at turbine
inlet air temperature levels which are beyond the structural temperature
--6--
ll~lSZ76
capabilities of high temperature alloys, turbine 36 must be cooled to assure
its structural integrity. It can, therefore, be appreciated that as the tempera-ture of the hot exhaust gases exiting combustor 34 is increased, an increased
percentage of cooling air is required to cool the turbine. Traditionally, the
source of the coolant for the turbine 36 has been air bled from the discharge
of compressor 20 which is routed to and through the turbine in a manner well
known in the art. The compressor discharge has been the logical choice for
the coolant flow since the pressure of the compressor discharge airflow
(referred to hereinafter as the "cooling air") is high enough to drive the cooling
air through the tortuous path associated with the turbine structure. However,
because the cooling air has had work performed on it by the compressor, its
temperature level has been increased. And, as compressor compression ratios
are increased, and as aircraft velocities increase, a corresponding rise in the
temperature of the cooling air is experienced. As a result, an increasingly
higher percentage of cooling flow is required to cool the turbine to acceptable
temperature levels. As mentioned earlier, this cooling air must bypass the
combustor and perhaps one or more turbine stages before being returned to
the propulsive cycle, thus giving rise to a performance penalty in proportion
to the amount of cooling air used. It thus becomes advantageous to reduce the
amount of cooling air required.
Referring now to Figure 3 there is depicted in graphical form
the change in turbine relative cooling flow rates and specific fuel consumption
as a function of the change in cooling air temperature for a typical gas turbofan
engine of the variety depicted in Figure 1. As an illustration, an estimate of
the cooling airflow and specific fuel consumption reductions that can be realized
by cooling the turbine blade cooling air of a two-stage core engine turbine of
--7--
5276
current design is shown in Figure 3. It may be observed from the
figure that in this particular application, reducing the cooling
air temperature by 250F results in a 50 percent reduction in the
required cooling airflow rate with a corresponding reduction of 1.1
percent inspecific fuel consumption. It is clear from this simpl-
istic that great benefits can be obtained by reducing the
temperature of the turbine cooling air.
The present invention contemplates the use of the relatively
cool fan bypass stream as a heat sink to cool the cooling air.
Referring again to Figure 1, it may be seen that the engine is
provided with a means for capturing a portion of the relatively
cool bypass flow such as, for example, shroud 44 which circumscri~es
a portion of the length of core engine 12 within the bypass duct
to define a flow passage 46 (perhaps in the form of an annulus)
therebetween. Disposed within this passage is a heat exchanger 54,
preferably of the cross-flow tubular type which is described in
greater particularity in U.S. Patent No. 4,020,150 - Thomas G.
Wakeman, issued on October 17, 1978. Turbine cooling air is bled
from the compressor discharge through ports 48 in the core engine
casing 50 at various circumferential locations and routed through
at least one conduit 52 to the heat exchanger 54. The bypass air
portion captured by shroud 44 enters a diffuser section 55 where
the dynamic head of the captured portion is largely recovered and
ducted through the heat exchanger where it absorbs heat from the
turbine cooling air. This bypass air portion is then returned to
the fan duct at the discharge 56 of passage 46. The cooling air
thus cooled is routed via conduit 58 to the high pressure turbine
36 through compressor rear frame struts 60 and thereafter to
the expander nozzle 62 of a type taught by U.S. Patent No.
3,565,545 issued to Melvin Bobo et al on February 23, 1971,
which is assigned to the assignee of the present
' ! F~
-- 8 --
276
invention. The cooled cooling air then travels via passageway 64 to turbine 36
where it is used to perform the cooling function in a manner well known in the
art.
In order to permit the efficient return of the heated bypass flow
5 portion back into bypass duct 26, its static pressure must be matched to the
static pressure in the bypass duct at location 56 where the bleed portion is
reintroduced Thus, the total pressure drop of the bled portion, including the
pressure drop through the diffuser section 55, heat exchanger S4 and flow
passage 46 must be limited to a value less than or equal to the dynamic head of
10 the remainder of the bypass flow stream at the location where the bled portion
is reintroduced into the fan duct.
If, as is the case in existing gas turbofan engines for which the
present invention may wish to be adapted, the space available for ducting
through the compressor rear frame struts 60 is limited, the configuration of
15 Figure 1 may be modified as in Figure 2 by reducing the cooling flow rate
ducted through heat exchanger 54 and increasing the magnitude of the cooling
air temperature reduction in direct proportion to the reduction in flow rate.
While this design approach reduces the size of the required ducting 58, it will
generally result in some increase in heat exchanger weight in order to increase
20 the effectiveness of the heat exchanger. In such an embodiment, an auxiliary
hot flow of cooling air is bled from the core engine through apertures 66 and
68 in core engine inner casing structure 70. This uncooled auxiliary bleed air
is mixed with the cooled cooling air exiting the downstream end 72 of conduit
58 ahead of the expander nozzle 62 to obtain the desired final cooling air
25 temperature. The resulting mixture is then utilized to cool the hot turbine
components as in Figure 1 and in accordance with well known turbine cooling
principle s .
_g _
11~5~76
It thus becomes clear from the foregoing descriptions that the
statcd objects of the present invention have been attained in the embodiments
as depicted and that an engine configured in accordance with the present
invention will have sigr,ificant performance benefits over prior art gas turbo-
5 fan engines In particular, reliance has been placed on the well-established
concept of utilizing compressor bleed air as a turbine blade coolant. However,
the amount of compressor bleed air required has been substantially reduced,
thereby enhancing overall cycle performance. Conversely, an increase in
blade life can be achieved by maintaining the original coolant flow rate but by
10 reducing the temperature of the coolant, with essentially no further degrada-
tion in engine performance. Furthermore, the present invention is readily
adaptable to existing gas turbofan engines in that the components may be
designed and placed in the engine in such a manner that they do not substan-
tially change the configuration or design of nearby existing structure.
15 Furthermore, the heat exchanger is of the air-to-air variety and is completely
independent of the need for highly volatile coolant fluids which characterize
prior art turbine cooling concepts.
It should become obvious to one skilled in the art that certain
changes can be made to the above-described invention without departing frorn
20 the broad inventive concepts thereof. For example, while the present inven-
tion contemplates the cooling of the turbine cooiant by placing it in heat
exchange relationship with the abundant supply of fan bypass air in the gas
turbofan engine, the particular configurahon of the heat exchanger may take
many forms, such as heat exchangers of the single or multiple-pass variety.
25 Furthermore, it may be desirable to extract the cooling air from the
compressor 20 at a location other than the compressor discharge. In addition,
5~
the present invention may be used to cool the cooling air required for any of
a number of high temperature turbine components and is not limited to cooling
the cooling air required for turbine blades and vanes. It is intended that the
appended claims cover all such variations in the present invention's broader
5 inventive concepts,
Having thus described the invention, what is claimed as novel
and desired to be secured by Letters Patent _is:
