United States Patent c191
`Stuart
`
`I lllll llllllll Ill lllll lllll lllll lllll lllll lllll lllll lllll llllll Ill lllll llll
`US005317877A
`5,317,877
`[11] Patent Number:
`Jun. 7, 1994
`[45] Date of Patent:
`
`[54]
`
`INTERCOOLED TURBINE BLADE
`COOLING AIR FEED SYSTEM
`Inventor: Alan R. Stuart, Hamilton, Ohio
`[75]
`[73] Assignee: General Electric Company,
`Cincinnati, Ohio
`[21] Appl. No.: 923,676
`Aug. 3, 1992
`[22] Filed:
`Int. CJ.s ................................................ F02C 7/16
`(51]
`[52] U.S. CI ....................................... 60/736; 60/39.75
`[58] Field of Search ...................... 60/39.07, 728, 736,
`60/39.75; 415/115, 116
`References Cited
`U.S. PATENT DOCUMENTS
`2,891,382 6/1959 Broffitt ............................... 60/39.66
`3,017,159 1/1962 Foster et al ...................... 253/39.15
`3,814,539 6/1974 Klompas ............................... 416/95
`4,136,516 1/1979 Corsmeier ............................. 60/736
`4,137,705 2/1979 Andersen et al. .................. 60/39.08
`4,187,675 2/1980 Wakeman ........................... 60/39.75
`4,550,561 11/1985 Coffinberry .......................... 601736
`4,645,415 2/1987 Hovan et al ........................ 415/115
`4,882,902 11/1989 Reigel et al. ....................... 60/39.75
`5,059,093 10/1991 Khalid et al. ...................... 60/39.07
`5,163,385 11/1992 Mazeaud et al. ................... 60/39.07
`
`[56]
`
`FOREIGN PATENT DOCUMENTS
`120826 5/1988 Japan ..................................... 60/736
`112631 4/1990 Japan ..................................... 60/736
`
`OTHER PUBLICATIONS
`Roth et al. How to Use Fuel as a Heat Sink Space &
`Aeronautics, Mar., 1960 pp. 56-60.
`Primary Examiner-Louis J. Casaregola
`Attorney, Agent, or Firm-Jerome C. Squillaro; John R.
`Rafter
`ABSTRACT
`[57]
`A gas turbine engine having a compressor and an air(cid:173)
`cooled turbine is provided with a cooling system for
`decreasing the temperature of the turbine cooling air. A
`heat exchanger, mounted on the compressor casing,
`receives a portion of the pressurized air which is bled
`from the compressor. A heat sink medium is pumped
`through the heat exchanger into heat exchange relation(cid:173)
`ship with the pressurized air, thereby cooling the air.
`The cooled air is then further pressurized and routed to
`and circulated through the turbine rotor blades to pro(cid:173)
`vide improved cooling thereof. The intercooling of the
`compressor bleed air permits a reduction in the quantity
`of compressor air required for turbine rotor blade cool(cid:173)
`ing or allows an increase in turbine entry temperature
`and thus provi!}es an improvement in engine power and
`performance. In the case where the heat sink medium is
`engine fuel, the heat extracted from the compressor
`bleed air is returned to the engine operating cycle in the
`form of heated engine fuel.
`
`16 Claims, 3 Drawing Sheets
`
`GE-1007.001
`
`

`
`34
`
`FUEL
`PUMP
`
`10
`
`/
`
`-33
`
`78
`
`36
`
`FUEL
`CONTROL
`
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`16
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`FIG. 1
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`GE-1007.002
`
`

`
`U.S. Patent
`
`June 7, 1994
`
`Sheet 2 of 3
`
`5,317,877
`
`~
`
`~68
`
`r I'- 74
`
`72
`
`FIG. 2
`
`FIG. 4
`
`GE-1007.003
`
`

`
`U.S. Patent
`
`June 7, 1994
`
`Sheet 3 of 3
`
`5,317,877
`
`34
`
`FUEL
`PUMP
`
`78
`
`36
`
`FUEL
`CONTROL
`
`58
`
`110
`
`FIG. 3.
`
`FIG. 5
`
`GE-1007.004
`
`

`
`1
`
`5,317,877
`
`2
`cycle. Thus, while turbine blade cooling has inherent
`advantages, it also has associated therewith certain in(cid:173)
`herent disadvantages which are functions of the quan(cid:173)
`tity of cooling air used in cooling the turbine rotor
`5 blades.
`It will, therefore, be appreciated that engine perfor(cid:173)
`mance can be increased by reducing the amount of
`cooling air required by the turbine rotor blades. One
`system for accomplishing this goal was disclosed in U.S.
`10 Pat. No. 4,137,705 to Andersen et al. In accordance
`with that teaching, an aircraft gas turbine engine is
`provided with a turbine wherein the rotor disk bears a
`plurality of hollow, air-cooled turbine blades. Cooling
`air is bled from the compressor portion of the engine
`and routed radially inwardly into a compact heat ex(cid:173)
`changer connected to and rotatable with the compres-
`sor. Heat which has been introduced into the cooling air
`through the compression process is extracted within the
`heat exchanger by engine lubricating oil which is routed
`through the heat exchanger and into heat exchange
`relationship with the cooling air. The cooled cooling air
`is then directed from the heat exchanger and through
`the turbine rotor blades to provide improved cooling
`thereof. The lubricating oil is that which performs the
`usual engine lubrication function so that an additional
`coolant need not be carried by the aircraft. Subse-
`quently, this oil is cooled by engine fuel or the fan by(cid:173)
`pass stream airflow (in a gas turbofan engine) in a sta-
`tionary heat exchanger relatively remote from the tur(cid:173)
`bine. The use of the fuel as the final heat sink results in
`a partially regenerative engine in that most of the heat
`removed from the compressed air is reintroduced into
`the engine cycle as heated engine fuel.
`
`INTERCOOLED TURBINE BLADE COOLING AIR
`FEED SYSTEM
`
`FIELD OF THE INVENTION
`This invention relates to gas turbines and, more par(cid:173)
`ticularly, to a concept for efficiently reducing the tem(cid:173)
`perature of air used to cool high-temperature turbine
`rotor blades.
`
`BACKGROUND OF THE INVENTION
`It is well understood that turbine inlet temperature is
`a major determinant of the specific power available
`from a gas turbine. However, current turbines are lim(cid:173)
`ited in inlet temperature by the physical properties of 15
`the materials used to construct the turbines. To permit
`turbines to operate at gas stream temperatures which
`are higher than the temperatures which conventional
`materials can normally tolerate, considerable effort has
`been devoted to the development of sophisticated meth- 20
`ods of turbine cooling.
`In early gas turbine engine designs, cooling of high(cid:173)
`temperature components was limited to transferring
`heat to lower-temperature parts by the method of con(cid:173)
`duction, and air-cooling technology was limited to pass- 25
`ing relatively cool air across the face of the turbine
`rotor disks. In order to take advantage of the potential
`performance improvements associated with higher tur(cid:173)
`bine inlet temperatures, modern turbine cooling tech(cid:173)
`nology utilizes air-cooled hollow turbine nozzle vanes 30
`and rotor blades to permit operation at inlet gas temper(cid:173)
`atures in excess of 2000° F. Various techniques have
`been developed to cool these hollow blades and vanes.
`These incorporate two basic forms of air cooling, used
`either singly or in combination, depending upon the 35
`level of gas temperatures encountered and the degree of
`sophistication permissible. These basic forms of air
`cooling are known as convection and film cooling.
`However, the benefits obtained from sophisticated
`air-cooling techniques are at least partially offset by the 40
`extraction of the necessary cooling air from the propul(cid:173)
`sive cycle. The conventional source of coolant for a
`high-pressure turbine is air which is bled off the com(cid:173)
`pressor portion of the gas turbine engine and is routed
`to the hollow interior of the turbine blades. The quest 45
`for thermal efficiency has caused an increase in the
`compressor delivery air temperature. However, the
`compressor air, having a temperature much less than
`that of the turbine flow path gas stream, absorbs heat
`from the turbine blades to maintain the blades at an 50
`acceptable temperature. When this heated cooling air
`leaves the turbine blades, perhaps as a coolant film, this
`heat energy is Jost to the propulsive cycle since the
`cooling air is normally mixed with the exhaust gases and
`ejected from an engine nozzle. More particularly, the 55
`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
`reintroduced into the flow path gas stream downstream
`of the turbine nozzle, it does not return its full measure 60
`of work to the cycle as it expands through the turbine.
`Additionally, the reintroduction of cooling air into the
`gas stream produces a loss in gas stream total pressure.
`This is a result of the momentum mixing losses associ(cid:173)
`ated with injecting a relatively low-pressure cooling air 65
`into a high-pressure gas stream. The greater the amount
`of cooling air which is routed through the turbine
`blades, the greater the losses become in the propulsive
`
`SUMMARY OF THE INVENTION
`It is an object of the present invention to improve
`upon the teaching of U.S. Pat. No. 4,137,705 by elimi(cid:173)
`nating the lubricating oil from the heat exchange cycle
`and performing the coolant heat exchange or heat rejec(cid:173)
`tion external to the engine, thus providing scope for
`efficient heat exchanger design, including modulation
`of coolant temperature and easy access for inspection
`and replacement. In particular, it is an object of the
`invention to provide direct or indirect heat exchange
`between cooling air bled from the high-pressure com(cid:173)
`pressor and the fuel for the engine.
`Another object of the invention is to enhance the
`efficiency and increase the power available from a core
`engine of predetermined size.
`A further object of the invention is to provide an
`aircraft gas turbine engine in which the turbine rotor
`blades are cooled to withstand the high-temperature
`gases of combustion.
`It is another object of the invention to reduce the
`amount of cooling air required by the turbine rotor
`blades by reducing the temperature of the cooling air
`passing therethrough in order to improve cooling effec(cid:173)
`tiveness.
`Yet another object is to provide an aircraft gas tur(cid:173)
`bine engine wherein the work done by the compressor
`on that portion of the pressurized air used for turbine
`cooling is returned to the engine power cycle as useful
`energy.
`Another object of the invention is to provide a mech(cid:173)
`anism for modulating the coolant temperature at lower
`power and lower fuel flows. In particular, to avoid
`overheating of the fuel at low fuel flows, the cooling air
`
`GE-1007.005
`
`

`
`5,317,877
`
`3
`bled from the compressor should bypass the fuel heat
`exchanger.
`These objects are attained in an aircraft gas turbine
`engine in accordance with the invention by providing a
`turbine wherein the rotor disk bears a plurality of hol- 5
`low, air-cooled turbine blades. Cooling air is bled from
`the high-pressure compressor and routed into a heat
`exchanger mounted at a location remote from the tur(cid:173)
`bine, for example, on the outside of the compressor
`casing. Heat which has been introduced into the cooling 10
`air through the compression process is extracted within
`the heat exchanger by fuel which acts as a heat sink
`when routed through the heat exchanger and brought
`into heat exchange relationship with the cooling air,
`either directly or by way of an intermediate inert or 15
`nonflammable fluid medium. The use of inert or non(cid:173)
`flammable fluid as the intermediate medium eliminates
`the possibility of fuel entering the cooling air flow path,
`which could result in the event of a leak in a heat ex(cid:173)
`changer where the fuel and cooling air are in direct heat 20
`exchange relationship.
`The cooled cooling air is then directed from the heat
`exchanger, further compressed and then passed through
`the turbine rotor blades to provide improved cooling
`thereof. For higher coolant flows, fan bypass air can be 25
`used as a supplemental or alternative heat sink. The use
`of fuel as the heat sink results in a partially regenerative
`engine in which the heat removed from the compressed
`air is reintroduced more efficiently into the engine cy(cid:173)
`cle.
`Incorporation of the turbine blade cooling system in
`accordance with the invention into an aircraft or other
`gas turbine engine permits a reduction in the quantity of
`compressor air required for turbine rotor blade cooling
`and thus provides an improvement in engine perfor- 35
`mance. Conversely, an increase in blade life can be
`achieved by maintaining the original coolant flow rate
`but reducing the temperature of the coolant, with essen(cid:173)
`tially no degradation in engine performance, or the
`turbine entry temperature can be increased to raise 40
`power output.
`
`4
`assembly (not shown) including a stage of fan blades
`(not shown), and a fan turbine (not shown) which is
`interconnected to the fan assembly by rotatable shaft 16.
`The core engine 12 includes an axial-flow high-pressure
`compressor 18 having a rotor 20 and a compressor
`casing 22 bearing a plurality of stators 24 interposed in
`alternating relationship with the stages of rotor 20.
`Each stage of rotor 20 bears a plurality of radially di(cid:173)
`rected, circumferentially distributed rotor blades 26 and
`each stator 24 bears a plurality of radially directed,
`circumferentially distributed stator guide vanes 28.
`Air enters the inlet (not shown) of and is initially
`compressed by the fan assembly. A first portion of this
`compressed air enters the fan bypass duct defined, in
`part, by core engine 12 and a circumscribing fan nacelle
`(not shown) and discharges through a fan duct 29 (only
`a portion of which is shown) and a fan nozzle (not
`shown). A second portion of the compressed air may be
`further compressed by a booster or other compressor
`and then enters inlet 30, is further compressed by the
`axial-flow compressor 18 and then is discharged to a
`combustor 32. In the combustor 32 the air is mixed with
`fuel. The fuel is supplied to fuel manifold 33 by means
`such as a fuel pump 34 and an engine fuel control 36 of
`a type well known in the art and responsive to pilot
`throttle inputs. The mixture is burned to provide high-
`energy combustion gases which drive a core engine
`turbine rotor 38.
`Core engine high-pressure turbine rotor 38 comprises
`30 a turbine disk 40 bearing a plurality of hollow turbine
`rotor blades 42 about its periphery. The turbine rotor 38
`drives, in tum, the compressor rotor 20 through inter(cid:173)
`connecting shaft 44 in the usual manner of a gas turbine
`engine. A stationary row of turbine nozzle vanes 46
`orients the flow into the rotating turbine rotor blades
`42.
`The hot combustion gases then pass through and
`drive the fan turbine, which in tum drives the fan as(cid:173)
`sembly. A propulsive force is thus obtained by the ac(cid:173)
`tion of the fan assembly discharging air from the fan
`bypass duct through the fan nozzle and by the discharge
`of combustion gases from a core engine nozzle (not
`shown), the structure of which is well known in the art.
`In accordance with a first preferred embodiment of
`the present invention, a turbine cooling system is pro(cid:173)
`vided which bleeds air from the high-pressure compres(cid:173)
`sor 18, transfers heat from that compressor bleed air to
`the fuel to be fed to the combustor 32, and then supplies
`the cooled compressor bleed air to the cooling circuits
`(not shown) of the rotor blades 42 of the high-pressure
`turbine rotor 38. The turbine cooling system generally
`includes an annular outlet manifold 48 for bleeding air
`from the high-pressure compressor 18, a heat exchanger
`50 for transferring heat by conduction from the com(cid:173)
`pressor bleed air to the fuel being fed to the combustor
`32, an annular inlet manifold 52 for circumferentially
`distributing the cooled compressor bleed air returned to
`the core engine from the heat exchanger, and an impel(cid:173)
`ler 54 for further compressing and feeding the cooling
`air to the hollow turbine rotor blades 42.
`In accordance with the invention, the compressor
`bleed air is extracted through a plurality of openings 56
`which communicate with outlet manifold 48. For the
`purpose of illustration only, FIG. shows the pressurized
`65 air being extracted behind the fourth-stage rotor, al-
`though the air may in the alternative be extracted fur-
`ther downstream or further upstream. The precise point
`of extraction will be a function of the amount of pressur-
`
`60
`
`BRIEF DESCRIPTION OF THE DRAWINGS
`These and other advantages of the invention will be
`better understood when the detailed description of the 45
`preferred embodiment of the invention is read in con(cid:173)
`junction with the drawings, wherein:
`FIG. 1 is a partial cross-sectional view of an aircraft
`gas turbine turbofan engine in accordance with a first
`preferred embodiment of the invention and illustrating 50
`schematically the relationship of various systems;
`FIG. 2 is an enlarged, fragmentary, cross-sectional
`view depicting the internal construction of the pin or fin
`heat exchanger in accordance with the invention;
`FIG. 3 is a block diagram illustrating schematically 55
`the relationship of various systems in accordance with a
`second preferred embodiment of the invention;
`FIG. 4 is a front view of the coolant impeller and
`turbine disk in accordance with the preferred embodi-
`ments of the invention; and
`FIG. 5 is a sectional view depicting the manner in
`which the coolant impeller of FIG. 4 is mounted.
`
`DETAILED DESCRIPTION OF THE
`PREFERRED EMBODIMENTS
`In FIG. an aircraft gas turbofan engine incorporating
`the invention is generally indicated by the numeral 10.
`This engine generally comprises a core engine 12, a fan
`
`GE-1007.006
`
`

`
`5,317,877
`
`s
`6
`In accordance with a third preferred embodiment of
`ization required in any particular gas turbine engine in
`the invention, fan bypass air is used as a secondary or
`conjunction with impeller 54 to force the cooling air
`alternative heat sink. Although the use of fan bypass air
`through blades 42.
`as a heat sink provides minimal regenerative benefit, it
`The high-pressure bleed air flows from outlet mani-
`fold 48 to heat exchanger 50 by way of outlet conduit 5 enables the compressor bleed air to be cooled, with
`58. Heat exchanger 50 comprises a casing 60 with an
`consequent reduction in the metal temperature of the
`inlet 62 for receiving pressurized bleed air from outlet
`turbine rotor blades, in cases where the fuel cannot
`conduit 58 and an outlet 64 for outputting cooled bleed
`serve as a heat sink for the compressor bleed air.
`air to an inlet conduit 66 for return to the core engine.
`A common feature of all preferred embodiments of
`A plurality of straight extruded tubes 68 are arranged 10 the invention is that the cooled bleed air exits the heat
`exchanger and is piped via inlet conduit 66 to an inlet 80
`inside casing 60 in a generally parallel array, the ends of
`straight tubes 68 being connected by a plurality of U-
`which communicates with annular inlet manifold 52 by
`shaped tubes 70 to form a serpentine heat exchange
`means such as a duct 82. The inlet manifold 52 circum-
`circuit.
`ferentially distributes the cooling air. From inlet mani-
`The internal structure of straight tube 68 is shown in 15 fold 52, the cooling air is then pumped into an annular
`detail in FIG. 2. Each straight tube 68 has associated
`cavity 84 between the compressor 18 and tube 85 by
`therewith a plurality of laterally extending pins or fins
`way of holes through stub shaft 86 in a well-known
`72 distributed at equal intervals along its length. The
`manner.
`fins are generally parallel, the parallel surfaces defining
`In the regenerative preferred embodiments, the 300°
`the direction of the compressor bleed air flowing there- 20 F. cooling air then flows aftward until it reaches an
`annular chamber 88 located inside the bore 92 of radial
`between from inlet 62 to outlet 64 of the heat ex-
`changer. Each pinned or finned tube 68 has an axially
`outflow impeller 54 and in front of (or to the rear of)
`extending hole 74 extending through the interior
`turbine disk 40. The pressure inside chamber 88 is about
`135 psi.
`thereof via which the fuel from fuel pump 34 flows on
`its way to fuel control 36.
`25 As best seen in FIG. 4, impeller 54 comprises a gener-
`In accordance with the first preferred embodiment of
`ally annular disk 94 having a plurality of hollow radial
`the invention, the inlet of the serpentine heat exchange
`spokes 96 circumferentially distributed on its periphery.
`circuit is connected to fuel tank 34 via fuel line 76; the
`Impeller 54 is seated on an arm 98 which extends from
`outlet of the serpentine heat exchange circuit is con-
`the high-pressure turbine disk 40. As best seen in FIG.
`nected to fuel control 36 via fuel line 78. The inlet and 30 5, flange 114 of shaft 44, flange 116 of arm 98 and flange
`118 of impeller 54 are secured to disk 120 by bolt 122.
`outlet of the serpentine heat exchange circuit are ar-
`ranged so that the fuel is in counterflow relationship
`During rotation of impeller 54, the cooling air in
`chamber 88 is centrifuged via radial holes 100, each of
`with the compressor bleed air.
`which extends from the bore 92 to the tip of a corre-
`By way of example, the compressor bleed air taken
`from the fourth stage of the high-pressure compressor 35 sponding spoke 96. Depending on conditions, impeller
`has a temperature of about 700° F. and a pressure of
`54 will have a pressure ratio of 2 or more. For example,
`about 150 psi. Inside the heat exchanger, that tempera-
`impeller 54 compresses the cooling air to a pressure of
`ture is reduced by the conduction of heat from the
`about 280 psi and a temperature of about 476° F. The
`compressor bleed air to the external surface of the
`compressed cooling air from impeller 54 then enters the
`finned tubes 68, further conduction of heat from the 40 spaces 102 formed between the roots 104 of the rotor
`external surface to the internal surface of finned tubes
`blades 106 and the corresponding dovetail slots 108
`68, and consequent conduction of heat from the internal
`formed in the turbine disk 40.
`surface of finned tubes 68 to the fuel. As a result of this
`Each rotor blade has a cooling circuit (not shown) of
`heat exchange, the temperature of the bleed air can be
`conventional design incorporated therein, which cool-
`reduced by up to 400° F., i.e., to a temperature of 300° 45 ing circuit communicates with the corresponding space
`102 via one or more inlets formed in the root portion
`F., while the temperature of the fuel is raised by 150' F.
`thereof. The rotor blade is then convection and film
`As a result, the heat removed from the compressor
`bleed air is recovered and returned to the engine pro-
`cooled by the cooling air which flows through the cool-
`pulsive cycle, thus improving overall engine perfor-
`ing circuit in a well-known manner.
`mance.
`The result of the intercooled cooling air system in
`accordance with the invention is a considerable reduc-
`In accordance with a second preferred embodiment
`of the invention, an inert or nonflammable fluid me-
`tion in the coolant flow and the coolant parasitic power
`dium, e.g., water or an antifreeze mixture such as water
`consumption, as compared to a conventional high-pres-
`and glycol, is placed in intermediate heat exchange
`sure turbine blade cooling air feed system.
`relationship for facilitating heat transfer from the com- 55
`In the regenerative preferred embodiments, the heat
`extracted from the compressor bleed air is not lost to
`pressor bleed air to the fuel. This preferred embodiment
`requires two heat exchangers: the first for the transfer of
`the cycle, but rather is added to the combustion process
`heat from the compressor bleed air to the intermediate
`via the fuel. This adds to the fuel energy input. Thus, an
`inert or nonflammable fluid and the second for the
`improvement in efficiency and power output can be
`transfer of heat from the intermediate inert or nonflam- 60 expected.
`The greatest benefit of the invention derives not so
`mable fluid to the fuel. The inert or nonflammable fluid
`would be pumped by pump 112 through a closed circuit
`much from the reduced coolant airflows, but rather
`which includes the serpentine heat exchange circuits of
`from the ability to considerably raise the turbine entry
`both heat exchangers 50 and 50', as depicted in FIG. 3.
`temperature, while reducing the metal temperature of
`The advantage of this construction is that in the event of 65 the high-pressure turbine rotor blades.
`In accordance with a further feature of the invention
`fuel leakage into the second heat exchanger, the fuel
`shown in FIG. 1, at low power states and fuel flow rates
`leakage will not enter the core engine with the cooling
`air, which would create a fire hazard.
`below a predetermined threshold, e.g., during cruise
`
`50
`
`GE-1007.007
`
`

`
`5,317,877
`
`7
`8
`cation with said heat exchange circuit of said first heat
`and idle, a bypass valve 110 can be operated to enable
`the coolant to bypass the heat exchanger, thereby mod-
`exchanger, thereby forming a closed circuit which
`carries said first fluid medium, whereby said second
`ulating the coolant temperature. This bypass feature
`avoids overheating of the fuel which is passing through
`fluid medium is heated in said second heat exchanger by
`the heat exchanger at a relatively low rate reflecting the S conduction of heat from said first fluid medium.
`low power state of the engine, which overheating could
`7. The gas turbine engine as defined in claim 6, fur-
`create fuel system problems.
`ther comprising a combustor and means for transport-
`The preferred embodiment has been described in
`ing said second fluid medium from said chamber outlet
`detail hereinabove for the purpose of illustration only.
`of said second heat exchanger to said combustor, said
`Various modifications could be made to the above- 10 first fluid medium being inert and said second fluid
`described structu~e wit~out departing. from th7 spirit
`medium being fuel.
`8. A system for feeding cooling air to the rotor blades
`and scope. of the mvent1on as defi1:1ed m the cl~s set
`f?rth heremafter .. For ex~ple, while the present mven·
`of a turbine in a gas turbine engine having a compressor
`t1on. has _bee1:1 depicted as mtegral part ?fag.as turbo-fan
`for compressing air and a combustor for burning a mix·
`engme, ~twill~ apparent to those ~killed_ m !he art of 15 ture of pressurized air and fuel, comprising:
`means for bleeding air from an intermediate stage of
`gas ~urbme engm~ that the present mve!1tlon 1~ equally
`applicable t~ engme~ of the gas turbojet vanety, ~as
`said compressor;
`turbofan engmes havmg three or more spools, or marme
`a first heat exchanger having a chamber with an inlet
`and. indu.st?al g~ turbines. For marine and ind~st~al
`and an outlet;
`engmes! 1t 1s possible to use water and ~tmosphe1:1c arr, 20 means for supplying said compressor bleed air from
`respect.1vely, as a secondary or alternative heat smk.
`an intermediate stage of compressor to said cham·
`I claim:
`b
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`er
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`rrst eat exc anger;
`1. A g~s tur~Hne engm~ compnsmg a co~press~r for
`means for circumferentially distributing said com-
`compre~smg arr a~d hav1:11g a rotor, a turbme havm¥ a
`pressor bleed air from said chamber outlet of said
`rotor with a plurality of air-cooled rotor blades, a casmg 25
`fi th d
`h
`t
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`d
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`h. h
`fi
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`rrs ea exc anger o an annu ar cav1 y surroun -
`w 1c encases sa1 compressor, a irst eat exc anger
`.
`h ft b
`hi h
`'d t b'
`·d
`t
`d ·
`mg a s a Y w c saJ ur me ro or nves saJ
`mounted outside said casing and having a chamber with
`compre~sor ro~or; a1:1d
`an inlet and an outlet, means for bleeding air from an
`.
`mean~ for tmpellmg said cool:d compres~or bleed air
`intermediate stage of said compressor, means for sup-
`radially outwar~, further mcr~mg its pressur~,
`plying said compressor bleed air to said chamber inlet of 30
`said first heat exchanger, and means for circumferen-
`through a plurality of hollow radial spokes to said
`roto.r bl~des,
`tially distributing said compressor bleed air from said
`.
`wherem S~J~ first h~at exch~ger ~urther h~ an mlet
`chamber outlet of said first heat exchanger to an annular
`for rece1vmg an mert fluid medium havm~ a tem-
`cavity surrounding a shaft by which said turbine rotor
`p~rature lower than the temper~ture of sai~ bl7ed
`drives said compressor rotor and means for impelling 35
`at~ and. a h~at exc~ang~ c1rcmt commll:mcatmg
`said cooled compressor bleed air radially outward, fur-
`with sa1? fluid me?1um m~et ~or cond_uctmg ~eat
`ther increasing its pressure, through a plurality of ho!-
`from said ?Ie.ed air ~o said. me~ fluid mediu~,
`low radial spokes to said air-cooled rotor blades, said
`whereby said mert ~u~d medium 1s heated and said
`first heat exchanger further having an inlet for receiving
`a first fluid inert medium and having a temperature 40
`compress?r bI:ed air ts cooled.
`.
`.
`9. T~e c?olmg air feed system as defme~ 11:1 claim ~·
`lower than the temperature of said bleed air and a heat
`whe~em said fn:st heat exchanger plac_es_sa1d mert fluid
`exchange circuit communicating with said fluid me-
`dium inlet for conducting heat from said bleed air to
`medium and said c<;>mpr7ssor bleed arr m counterflow
`said first fluid medium, whereby said first fluid medium
`heat exchange relat10nsh1p.
`is heated and said compressor bleed air is cooled.
`10. The cooling air feed system as defmed in claim 8,
`2. The gas turbine engine as defined in claim 1,
`wherein said means for supplying said compressor bleed
`wherein said first fluid medium is water.
`air to said chamber inlet of said first heat exchanger
`3. The gas turbine engine as defmed in claim 1,
`comprises an annular manifold in fluid communication
`wherein said frrst fluid medium is air.
`with said compressor.
`11. The cooling air feed system as defmed in claim 8,
`4. The gas turbine engine as defined in claim 1, 50
`wherein said frrst heat exchanger places said frrst fluid
`further comprising a second heat exchanger having a
`medium and said compressor bleed air in counterflow
`chamber with an inlet and an outlet, means for supply-
`heat exchange relationship.
`ing fuel to said chamber inlet of said second heat ex-
`5. The gas turbine engine as defined in claim 1,
`changer, and means for transporting said fuel from said
`wherein said means for supplying said compressor bleed 55 chamber outlet of said second heat exchanger to said
`air to said chamber inlet of said first heat exchanger
`combustor, wherein said fuel has a temperature lower
`comprises an annular manifold in fluid communication
`than the temperature of said inert fluid medium and said
`with said compressor via a plurality of outlets formed in
`second heat exchanger further has a heat exchange
`said casing.
`circuit in fluid communication with said heat exchange
`6. The gas turbine engine as defined, in claim 1, fur- 60 circuit of said first heat exchanger, thereby forming a
`ther comprising a second heat exchanger mounted out-
`closed circuit which carries said inert fluid medium,
`side said engine casing and having a chamber with an
`whereby said fuel is heated before being transported to
`inlet and an outlet, and means for supplying a second
`said combustor by conduction of heat from said inert
`fluid medium to said chamber inlet of said second heat
`fluid medium.
`12. The cooling air feed system as defmed in claim 8,
`exchanger, wherein said second fluid medium has a 65
`temperature lower than the temperature of said first
`wherein said fluid medium is water.
`fluid medium and said second heat exchanger further
`13. The cooling air feed system as defined in claim 8,
`has a heat exchange circuit which is in fluid communi-
`wherein said fluid medium is antifreeze.
`
`45
`
`GE-1007.008
`
`

`
`5,317,877
`
`9
`14. A system for feeding cooling air to the rotor
`blades of a turbine in a gas turbine engine having a
`compressor for compressing air and a combustor for
`burning a mixture of pressurized air and fuel, compris- 5
`ing:
`means for bleeding air from an intermediate stage of
`said compressor;
`heat exchange means for transferring heat by conduc(cid:173)
`tion from a first fluid communication circuit to a lO
`second fluid communication circuit;
`means for supplying said compressor bleed air from
`an intermediate stage of said compressor to said
`first fluid communication circuit of said heat ex(cid:173)
`change means;
`means for supplying fuel having a temperature less
`than the temperature of said compressor bleed air
`to said second fluid communication circuit of said
`heat exc