(19) TZZ¥_Z__T

(11) EP 3 170 981 A1

(12) EUROPEAN PATENT APPLICATION

(43) Date of publication: (51) Int Cl.: 24.05.2017 Bulletin 2017/21 F01D 5/18 (2006.01)

(21) Application number: 16200307.3

(22) Date of filing: 23.11.2016

(84) Designated Contracting States: (71) Applicant: United Technologies Corporation AL AT BE BG CH CY CZ DE DK EE ES FI FR GB Farmington, CT 06032 (US) GR HR HU IE IS IT LI LT LU LV MC MK MT NL NO PL PT RO RS SE SI SK SM TR (72) Inventor: THORNTON, Lane Designated Extension States: Tolland, CT Connecticut 06084 (US) BA ME Designated Validation States: (74) Representative: Leckey, David Herbert MA MD Dehns St Bride’s House (30) Priority: 23.11.2015 US 201514948511 10 Salisbury Square London EC4Y 8JD (GB)

(54) BAFFLE FOR A COMPONENT OF A GAS TURBINE ENGINE

(57) An airfoil (60;160;260) according to an example (76) for conveying coolant (F). The baffle body (74) is of the present disclosure includes, an airfoil body (61) situated in the cavity (64,64A-C) such that a majority of defining a cavity (64,64A-C), and a external baffle surfaces of the sidewalls (78) abut the cavity (72,72’,72",72"’;172;272) including a baffle body (74) in- (64,64A-C). cluding sidewalls (78) and defining an internal passage EP 3 170 981 A1

Printed by Jouve, 75001 PARIS (FR) 1 EP 3 170 981 A1 2

Description rial. [0013] In a further embodiment of any of the foregoing BACKGROUND embodiments, the airfoil body extends between a plat- form and an airfoil tip. The cavity extends inwardly from [0001] This disclosure relates to cooling for a compo- 5 the airfoil tip, and the baffle body is situated in the cavity nent of a gas turbine engine. such that the internal passage is configured to eject cool- [0002] Gas turbine engines can include a fan for pro- ant adjacent the airfoil tip. pulsion air and to cool components. The fan also delivers [0014] A gas turbine engine according to an example air into a core engine where it is compressed. The com- of the present disclosure, and which the Applicant ex- pressed air is then delivered into a combustion section, 10 pressly reserves the right to claim, includes a rotor where it is mixed with fuel and ignited. The combustion spaced axially from a vane. At least one of the rotor and gas expands downstream over and drives turbine blades. the vane includes an airfoil body. The airfoil body includes Static vanes are positioned adjacent to the turbine blades external walls extending between a leading edge and a to control the flow of the products of combustion. The trailing edge, the external walls defining a cavity, and a blades and vanes are subject to extreme heat, and thus 15 baffle including a baffle body defining an internal passage cooling schemes are utilized for each. for conveying coolant. Sidewalls of the baffle body have a complementary geometry with the cavity. SUMMARY [0015] In an embodiment of the foregoing embodi- ment, sidewalls of the baffle body abut a majority of sur- [0003] An airfoil according to an example of the present 20 faces of the cavity. disclosure includes an airfoil body defining a cavity, and [0016] In a further embodiment of any of the foregoing a baffle including a baffle body including sidewalls and embodiments, the baffle body includes an inlet region defining an internal passage for conveying coolant. The and an exit region. The inlet region is configured to re- baffle body is situated in the cavity such that a majority ceive coolant, and the sidewalls are spaced apart at the of external surfaces of the sidewalls abut the cavity. 25 exit region to define one or more exit ports configured to [0004] In an embodiment of the foregoing embodi- eject coolant adjacent to an external surface of the airfoil ment, the sidewalls define an intermediate region be- body. tween an inlet region and an exit region. The inlet region [0017] In a further embodiment of any of the foregoing is configured to receive coolant, and the sidewalls are embodiments, the sidewalls taper from the inlet region spaced apart at the exit region to define one or more exit 30 towards the exit region. ports configured to eject coolant outwardly of the cavity. [0018] In a further embodiment of any of the foregoing [0005] In a further embodiment of any of the foregoing embodiments, the baffle body defines one or more exit embodiments, the inlet region tapers towards the inter- ports configured to eject coolant outward of the cavity. mediate region. [0019] In a further embodiment of any of the foregoing [0006] In a further embodiment of any of the foregoing 35 embodiments, the airfoil body is made of a first material, embodiments, the airfoil body extends in a chordwise and the baffle body is made of a second, different material direction between a leading edge and a trailing edge, and having a lesser thermal resistance than the first material. at least some of the exit ports are situated adjacent to [0020] A method of repairing an airfoil according to an the trailing edge. example of the present disclosure includes providing an [0007] In a further embodiment of any of the foregoing 40 airfoil body. The airfoil body has external walls extending embodiments, the exit region of the baffle body extends between a leading edge and a trailing edge providing a in the chordwise direction outwardly of the trailing edge. baffle. The baffle includes a baffle body defining an in- [0008] In a further embodiment of any of the foregoing ternal passage. Sidewalls of the baffle body define a first embodiments, the cavity is bounded by external walls of contour defining a cavity. The cavity extends inwardly the airfoil body. 45 from the external walls to define a second contour com- [0009] In a further embodiment of any of the foregoing plementary to the first contour. The method includes in- embodiments, the sidewalls of the baffle body have a serting the baffle into the cavity. complementary geometry to the surfaces of the cavity. [0021] In an embodiment of the foregoing embodi- [0010] In a further embodiment of any of the foregoing ment, the step of defining the cavity includes removing embodiments, the baffle body includes a plurality of cool- 50 material from the trailing edge to define an opening to ing features within the internal passage. the cavity, and the sidewalls of the baffle body are spaced [0011] In a further embodiment of any of the foregoing apart by an exit wall to define one or more exit ports embodiments, at least some of the plurality of cooling situated adjacent to the opening. features extend between opposed surfaces of the inter- [0022] In a further embodiment of any of the foregoing nal passage. 55 embodiments, the airfoil body is made of a first material, [0012] In a further embodiment of any of the foregoing and the baffle body is made of a second, different mate- embodiments, the airfoil body is made of a first material, rial. and the baffle body is made of a second, different mate- [0023] Although the different examples have the spe-

2 3 EP 3 170 981 A1 4 cific components shown in the illustrations, embodiments [0028] The engine 20 generally includes a low speed of this disclosure are not limited to those particular com- spool 30 and a high speed spool 32 mounted for rotation binations. It is possible to use some of the components about an engine central longitudinal axis A relative to an or features from one of the examples in combination with engine static structure 36 via several bearing systems features or components from another one of the exam- 5 38. It should be understood that various bearing systems ples. 38 at various locations may alternatively or additionally [0024] The various features and advantages of this in- be provided, and the location of bearing systems 38 may vention will become apparent to those skilled in the art be varied as appropriate to the application. from the following detailed description of an embodiment. [0029] The low speed spool 30 generally includes an The drawings that accompany the detailed description 10 inner shaft 40 that interconnects a fan 42, a first (or low) can be briefly described as follows. pressure compressor 44 and a second (or low) pressure turbine 46. The inner shaft 40 is connected to the fan 42 BRIEF DESCRIPTION OF THE DRAWINGS through a speed change mechanism, which in exemplary gas turbine engine 20 is illustrated as a geared architec- [0025] 15 ture 48, to drive the fan 42 at a lower speed than the low speed spool 30. Figure 1 schematically shows a gas turbine engine. [0030] The high speed spool 32 includes an outer shaft Figure 2 schematically shows an airfoil arrangement 50 that interconnects a second (or high) pressure com- for a turbine section. pressor 52 and a first (or high) pressure turbine 54. A Figure 3A illustrates a perspective view of an airfoil 20 combustor 56 is arranged between the high pressure and a baffle. compressor 52 and the high pressure turbine 54. A mid- Figure 3B illustrates a cross-sectional view of the turbine frame 57 of the engine static structure 36 is ar- airfoil of Figure 3A taken along line 3B-3B. ranged generally between the high pressure turbine 54 Figure 3C illustrates an isolated perspective view of and the low pressure turbine 46. The mid-turbine frame the baffle of Figure 3A. 25 57 further supports the bearing systems 38 in the turbine Figure 3D illustrates a sectioned, axial view of se- section 28. The inner shaft 40 and the outer shaft 50 are lected portions of the baffle of Figure 3C taken along concentric and rotate via bearing systems 38 about the line 3D-3D. engine central longitudinal axis A, which is collinear with Figure 4A illustrates an airfoil and a baffle according their longitudinal axes. to a second example. 30 [0031] The core airflow is compressed by the low pres- Figure 4B illustrates an airfoil and a baffle according sure compressor 44 then the high pressure compressor to a third example. 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pres- DETAILED DESCRIPTION sure turbine 46. The mid-turbine frame 57 includes air- 35 foils 59 which are in the core airflow path C. The turbines [0026] Figure 1 schematically illustrates a gas turbine 46, 54 rotationally drive the respective low speed spool engine 20. The gas turbine engine 20 is disclosed herein 30 and high speed spool 32 in response to the expansion. as a two-spool turbofan that generally incorporates a fan It will be appreciated that each of the positions of the fan section 22, a compressor section 24, a combustor section section 22, compressor section 24, combustor section 26 and a turbine section 28. The concepts described40 26, turbine section 28, and fan drive gear system 48 may herein are not limited to use with turbofans and may be be varied. For example, gear system 48 may be located applied to other types of turbine engines, such as three- aft of combustor section 26 or even aft of turbine section spool architectures. Alternative engines might also in- 28, and fan section 22 may be positioned forward or aft clude an augmentor section (not shown) among other of the location of gear system 48. systems or features, or, may not include the fan section 45 [0032] The engine 20 in one example is a high-bypass 22, such as in industrial gas turbine engines. geared aircraft engine. In a further example, the engine [0027] The fan section 22 drives air along a bypass 20 bypass ratio is greater than about six, with an example flow path B in a bypass duct defined within a nacelle 15, embodiment being greater than about ten, the geared while the compressor section 24 drives air along a core architecture 48 is an epicyclic gear train, such as a plan- flow path C for compression and communication into the 50 etary gear system or other gear system, with a gear re- combustor section 26 then expansion through the turbine duction ratio of greater than about 2.3 and the low pres- section 28. Although depicted as a two-spool turbofan sure turbine 46 has a pressure ratio that is greater than gas turbine engine in the disclosed non-limiting embod- about five. In one disclosed embodiment, the engine 20 iment, the examples herein are not limited to use with bypass ratio is greater than about ten, the fan diameter two-spool turbofans and may be applied to other types 55 is significantly larger than that of the low pressure com- of turbomachinery, including direct drive engine architec- pressor 44, and the low pressure turbine 46 has a pres- tures, three-spool engine architectures, and ground- sure ratio that is greater than about five. Low pressure based turbines. turbine 46 pressure ratio is pressure measured prior to

3 5 EP 3 170 981 A1 6 inlet of low pressure turbine 46 as related to the pressure one or more internal passages or cavities 64 that serve at the outlet of the low pressure turbine 46 prior to an to convey a fluid flow F from a coolant source 69 through exhaust nozzle. The geared architecture 48 may be an the airfoil 60. For example, the coolant source 69 can be epicycle gear train, such as a planetary gear system or relatively cool air from the compressor section 24, an other gear system, with a gear reduction ratio of greater 5 upstream stage of the turbine section 28, or bypass flow than about 2.3:1. It should be understood, however, that B from the fan section 22. Although not limited, the inter- the above parameters are only exemplary of one embod- nal cavities 64 in this example are defined by one or more iment of a geared architecture engine and that the internal walls or ribs 70. The fluid flow F is thus conveyed present invention is applicable to other gas turbine en- through the internal cavities 64 and is then discharged gines, including direct drive turbofans. 10 into the core flow path C through holes or openings in [0033] A significant amount of thrust is provided by the the airfoil body 61. In alternative examples, the airfoil bypass flow B due to the high bypass ratio. The fan sec- body61 defines a serpentine passage or cavity that winds tion 22 of the engine 20 is designed for a particular flight radially back and forth within the airfoil body 61 with one condition -- typically cruise at about 0.8 Mach and about or more ribs separating the passage sections. Although 35,000 feet (10,668 m). The flight condition of 0.8 Mach 15 the internal cavities 64 are depicted as extending in the and 35,000 ft (10,668 m), with the engine at its best fuel radial direction R, it should be appreciated that the inter- consumption - also known as "bucket cruise Thrust Spe- nal cavities can be arranged at different orientations rel- cific Fuel Consumption (’TSFC ’)" - is the industry stand- ative to each other and/or relative to the engine axis A ard parameter of lbm of fuel being burned divided by lbf to provide the desired cooling augmentation. of thrust the engine produces at that minimum point. "Low 20 [0036] Referring to Figures 3A-3D, a baffle 72 is situ- fan pressure ratio" is the pressure ratio across the fan ated in one of the internal cavities 64 of the airfoil 60. blade alone, without a Fan Exit Guide Vane ("FEGV") Figure 3A illustrates a perspective view of the airfoil 60. system. The low fan pressure ratio as disclosed herein Figure 3B illustrates a sectioned, radially inward view of according to one non-limiting embodiment is less than the airfoil 60 taken along line 3B-3B. Figure 3C illustrates about 1.45. "Low corrected fan tip speed" is the actual 25 an isolated perspective view of the baffle 72. Figure 3D fan tip speed in ft/sec divided by an industry standard illustrates a sectioned, axial view of the baffle 72 taken temperature correction of [(Tram °R) / (518.7 °R)]0.5 along line 3D-3D. In the illustrated example, the airfoil 9 (wherein °R = K x /5). The "Low corrected fan tip speed" body 61 defines internal cavities 64A-64C which can be as disclosed herein according to one non-limiting em- connected by one or more crossover passages 71 bodiment is less than about 1150 ft / second (350.5 m/s). 30 (shown in dashed lines) to convey fluid flow between one [0034] Figure 2 illustrates a portion of the turbine sec- or more of the internal cavities 64A-64C. Although three tion 28, such as one of the high or low pressure turbines internal cavities 64A-64C are shown, the airfoil 60 can 46, 54, which includes an airfoil 60. In this disclosure, define fewer or more than three internal cavities 64 and like reference numerals designate like elements where at various locations of the airfoil 60. The internal cavities appropriate and reference numerals with the addition of 35 64A-64C are bounded by external walls 68 and internal one-hundred or multiples thereof designate modified el- walls 70 of the airfoil body 61. ements that are understood to incorporate the same fea- [0037] The baffle 72 is situated in the internal cavity tures and benefits of the corresponding original ele- 64C adjacent to the trailing edge 67 of the airfoil 60. The ments. In this example, the airfoil 60 is a vane that is baffle 72 can be situated in other locations of the airfoil situated between two rotatable blades 62. Blade outer 40 60, such as in intermediate internal cavity 64B between air seal (BOAS) 58 is spaced radially outward from tip 73 pressure and suction sides P, S of the airfoil 60, and/or of the blade 62 to define a clearance gap G and to bound forward internal cavity 64A adjacent to the leading edge a portion of the core flow path C. The turbine section 28 65. In some examples, baffles 72’, 72" are situated ad- includes multiple airfoils or vanes 60, blades 62, and jacent to airfoil tip 73 or leading edge of blade 62 (Figure blade outer air seals 58 arranged circumferentially about 45 2). In one example, the baffle 72 extends at least partially the engine axis A. Although the examples herein are de- through one or more of the platforms 63 (Figure 2). In scribed with respect to the airfoil 60 as a vane, the ex- another example, baffle 72" is situated along or otherwise amples are also applicable to rotatable blades 62 and adjacent to tip 73 of blade 62 to eject coolant into the airfoils in other sections of the turbine section 28 or the clearance gap G (Figure 2). In one example, baffle 72"’ compressor section 24. Other portions of the engine 20 50 extends at least partially through BOAS 58 (Fig. 2), and may benefit from the teachings herein, such as combus- can be configured to eject coolant adjacent a trailing edge tor panels in the combustor section 26. face or mate face of BOAS 58, for example. [0035] The airfoil 60 includes an airfoil body 61 extend- [0038] The baffle 72 is configured to occupy a volume ingin a radial direction R between platforms 63.The airfoil of the internal cavity 64 to provide a desired cooling aug- body 61 extends in a chordwise direction C between a 55 mentation to portions of the airfoil 60 adjacent to the baffle leading edge 65 and a trailing edge 67, and in a circum- 72 or other heat loads. The baffle 72 includes an elon- ferential or thickness direction T between pressure and gated baffle body 74 extending between ends 77 (Figure suction sides P, S (Figure 3B). The airfoil body 61 defines 3C) and is sized to be received or situated in the internal

4 7 EP 3 170 981 A1 8 cavity 64C. A cross-section of the internal cavity 64C area for convective cooling and/or direct or meter fluid taken parallel to plane T, C defines a first contour, and flow within or through localized regions of the internal the baffle body 74 defines a second contour complemen- passage 76. Various cooling features 80 can include ped- tary to the first contour. In this arrangement, the baffle estals 80A and ribs 80B extending between opposed sur- body 74 has a complementary geometry to surfaces of 5 faces of the internal passage 76, for example. Other cool- the internal cavity 64C. The first and second contours ing features 80 can include features having a curved or can be taken as cross-sections parallel to plane T, C and complex geometry 80C (shown in dashed line) to direct along one or more intervals parallel to the radial axis R, flow through the internal passage 76, and trip strips 80D for example. In some examples, the first and second con- (shown in dashed line) or dimples protruding from sur- tours are substantially equal or equal at one or more of 10 faces of the internal passage 76 to cause turbulence in the ends 77 (shown in Figure 3B), at a majority of the the fluid flow F. intervals, or at each position between ends 77. In the [0043] The exit region 76C includes one or more exit illustrated example, a cross-section of the baffle body 74 ports 82 configured to eject fluid flow F outwardly of the is substantially constant between ends 77. internal passage 76. In the illustrated example, sidewalls [0039] In some examples, a volume of the baffle body 15 78 are spaced apart at the exit region 76C by exit wall 74 is substantially equal or equal to a volume of the in- 86 to define one or more exit ports 82 between ribs 80B ternal cavity 64C. The baffle body 74 is situated in the situated at the trailing edge 67. The ribs 80B can be ar- internal cavity 64C such that a majority or substantially ranged such that the exit ports 82 eject fluid flow F from each external surface of sidewalls 78 of the baffle body exit wall 86 at a desired orientation and/or velocity. In 74 abuts or directly contacts adjacent surfaces of the 20 alternative examples, the exit ports 82 are configured to cavity 64C. In this arrangement, fluid flow F through the eject fluid flow adjacent to another external surface of internal passage 76 provides convective cooling to por- the airfoil body 61, such as surfaces of the pressure or tions of the airfoil 60 adjacent to the internal cavity 64C. suction sides P, S of airfoil body 61 or blade tip 73. [0040] The baffle body 74 defines at least one internal [0044] The exit region 76C can be arranged at a de- passage 76 for conveying fluid flow F. The internal pas- 25 sired location relative to external surfaces of the airfoil sage 76 is bounded by sidewalls 78 of the baffle body 60 such as the trailing edge 67. For example, the exit 74. In the illustrated example, the sidewalls 78 define region 76C can be spaced a distance d1 in the chordwise inlet region 76A, intermediate region 76B and exit region direction C inwardly of the trailing edge 67 (Figure 3B). 76C of the internal passage 76. The inlet region 76A can In alternative examples, exit region 176C extend a dis- be provided with one or more inlet ports 79 for receiving 30 tance d2 in the chordwise direction C outwardly of the fluid flow F from the coolant source 69. The inlet region trailing edge 167 (Figure 4A) such that baffle 172 defines 76A can be configured to communicate fluid flow F from an axially aftmost portion of the trailing edge 167 and is the inlet ports 79 to the intermediate region 76B. The inlet exposed to the core flow path C. In another example, ports 79 can be defined at one or more of the ends 77 or surfaces of exit wall 86 of the exit region 276C are sub- sidewalls 78 (Figure 3D). In alternative examples, some 35 stantially flush with the trailing edge 267 (Figure 4B). By of the inlet ports 79 receive fluid flow, and other inlet ports arranging the exit region 76C/176C/276C relative to the 79 feed another cavity 64 with fluid flow, such as cross- trailing edge 67/167/276, a relative circumferential dis- over passage 71 coupling internal cavities 64B and 64C tance between adjacent airfoils 60 can be selected to to provide a desired relative cooling augmentation. In define a desired area of the airfoil stage. some examples, fluid flow is fed from openings at ends 40 [0045] The internal cavity 64 of airfoil 60 can be cast, 77 of the baffle body 74 to another cavity 64, such as machined or formed by an additive manufacturing tech- along the inlet or intermediate regions 76A, 76B. nique, for example. The baffle 72 can be situated in the [0041] The sidewalls 78 at the inlet region 76A can internal cavity 64 utilizing a casting or additive manufac- taper towards the intermediate region 76B to provide a turing technique, for example. In other examples, the baf- desired pressure differential between regions 76A and 45 fle 72 is situated in the internal cavity 64 subsequent to 76B or a desired surface area, for example. The tapered fabrication of the airfoil body 61. The baffle 72 can be arrangement can also be utilized to retain the baffle 72 utilized to form a sleeve or insert within the internal cavity within the internal cavity 64C or otherwise limit relative 64 such that different baffle configurations can be utilized movement of the baffle 72 and the airfoil 60. In examples, with a common predefined contour of the internal cavity the internal passage 76 between the inlet region 76A and 50 64. In this manner, the material and/or cooling charac- the exit region 76C is fluidly isolated from the internal teristics of the baffle 72 can be modified without having cavity 64C. In other examples, sidewalls 78 define one to modify a geometry of the airfoil 60. The techniques or more impingement cooling holes 84 (shown in dashed described herein can reduce fabrication cost and com- lines in Figure 3C) to provide impingement cooling to plexity by reducing casting die or casting core rework in adjacent surfaces of the internal cavity 64C. 55 an investment casting process, for example. The tech- [0042] The baffle 72 can include one or more cooling niques described herein can also improve cycle time dur- features 80 within the internal passage 76. The cooling ing iterations of airfoil redesign and retrofit. features 80 can be situated to provide additional surface [0046] In example repair techniques, material is re-

5 9 EP 3 170 981 A1 10 moved from the airfoil body 61 to define the internal cavity cavity(64,64A-C) suchthat a majority of external 64. Material can be removed from external surfaces of surfaces of the sidewalls (78) abut the cavity the airfoil body 61 to define an opening 81 to the internal (64,64A-C). cavity 64, such as at the trailing edge 67 of airfoil 60 with exit ports 82 situated at the opening 81 (Figures 3A-3B), 5 2. The airfoil as recited in claim 1, wherein the sidewalls for example. In some examples, the airfoil 60 or airfoil (78) define an intermediate region (76B) between an body 61 is made of a first material, and the baffle body inlet region (76A) and an exit region 74 is made of a second, different material. The first ma- (76C;176C;276C), the inlet region (76A) being con- terial of the airfoil 60 can be a high temperature material figured to receive coolant (F), and the sidewalls (78) such as a nickel based alloy cast as a single crystal, for 10 are spaced apart at the exit region (76C;176C;276C) example. The second material of the baffle body 74 can to define one or more exit ports (82;182;282) config- be selected to have a lesser thermal resistance than the ured to eject coolant (F) outwardly of the cavity first material, or which may be relatively lower cost or (64,64A-C). weight, for example. In some examples, the baffle 72 is fabricated of a metal or metal alloy, such as sheet metal, 15 3. The airfoil as recited in claim 2, wherein the inlet a multiple crystal nickel alloy or cobalt based alloy, region (76A) tapers towards the intermediate region formed by additive manufacturing, by casting, or the like. (76B). During installation, the baffle 72 is moved into the internal cavity 64C, such as from a radially outward to a radially 4. The airfoil as recited in claim 2 or 3, wherein the inward direction relative to axis R. The baffle 72 can be 20 airfoil body (61) extends in a chordwise direction (C) sized to form an interference fit with the internal cavity between a leading edge (65) and a trailing edge 64C. In alternative examples, the baffle 72 is fixedly at- (67;167;267), and at least some of the exit ports tached to the airfoil 60 by welding, fasteners, or the like. (82;182;282) are situated adjacent to the trailing In some examples, the baffle 72 is removable from the edge (67;167;267). airfoil 60. 25 [0047] Although particular step sequences are shown, 5. The airfoil as recited in claim 4, wherein the exit re- described, and claimed, it should be understood that gion (76C;176C;276C) of the baffle body (74) ex- steps may be performed in any order, separated or com- tends in the chordwise direction (C) outwardly of the bined unless otherwise indicated and will still benefit from trailing edge (67;167;267). the present disclosure. 30 [0048] It should be understood that relative positional 6. The airfoil as recited in any preceding claim, wherein terms such as "forward," "aft," "upper," "lower," "above," the cavity (64,64A-C) is bounded by external walls "below," and the like are with reference to the normal (68;168;268) of the airfoil body (61), wherein, option- operational attitude of the vehicle and should not be con- ally, the sidewalls (78) of the baffle body (74) have sidered otherwise limiting. 35 a complementary geometry to the surfaces of the [0049] The foregoing description is exemplary rather cavity (64,64A-C). thandefined by the limitationswithin. Variousnon-limiting embodiments are disclosed herein, however, one of or- 7. The airfoil as recited in any preceding claim, wherein dinary skill in the art would recognize that various mod- the baffle body (74) includes a plurality of cooling ifications and variations in light of the above teachings 40 features (80,80A-D;180;280) within the internal pas- will fall within the scope of the appended claims. It is sage (76), wherein, optionally, at least some of the therefore to be understood that within the scope of the plurality of cooling features (80...280) extend be- appended claims, the disclosure may be practiced other tween opposed surfaces of the internal passage than as specifically described. For that reason the ap- (76). pended claims should be studied to determine true scope 45 and content. 8. The airfoil as recited in any preceding claim, wherein the airfoil body (61) extends between a platform (63) and an airfoil tip (73), the cavity (64,64A-C) extends Claims inwardly from the airfoil tip (73), and the baffle body 50 (74) is situated in the cavity (64,64A-C) such that the 1. An airfoil (60;160;260), comprising: internal passage (76) is configured to eject coolant (F) adjacent the airfoil tip (73). an airfoil body (61) defining a cavity (64,64A-C); and 9. A gas turbine engine (20), comprising: a baffle (72,72’,72",72"’;172;272) including a55 baffle body (74) including sidewalls (78) and de- a rotor spaced axially from a vane, fining an internal passage (76) for conveying wherein at least one of the rotor and the vane coolant (F), the baffle body (74) situated in the includes an airfoil (60;160;260) of any preceding

6 11 EP 3 170 981 A1 12

claim, the airfoil body (61) including external adjacent to the opening (81). walls (68;168;268) extending between a leading edge (65) and a trailing edge (67;167;267), the 15. The method as recited in claim 13 or 14, wherein the external walls (68;168;268) defining the cavity airfoil body (61) is made of a first material, and the (64,64A-C); and 5 baffle body (74) is made of a second, different ma- the sidewalls (78) of the baffle body (74) having terial. a complementary geometry with the cavity (64,64A-C).

10. The gas turbine engine as recited in claim 9, wherein 10 sidewalls (78) of the baffle body (74) abut a majority of surfaces of the cavity (64,64A-C).

11. The gas turbine engine as recited in claim 9 or 10, wherein the baffle body (74) includes an inlet region 15 (76A) and an exit region (76C;176C;276C), the inlet region (76A) being configured to receive coolant (F), and the sidewalls (78) are spaced apart at the exit region (76C;176C;276C) to define one or more exit ports (82;182;282) configured to eject coolant (F) ad- 20 jacent to an external surface of the airfoil body (61), wherein, optionally, the sidewalls (78) taper from the inlet region (76A) towards the exit region (76C;176C;276C). 25 12. The gas turbine engine as recited in claim 9, 10 or 11, wherein the airfoil body (61) is made of a first material, and the baffle body (74) is made of a sec- ond, different material having a lesser thermal resist- ance than the first material. 30

13. A method of repairing an airfoil (60;160;260), com- prising:

providingan airfoil body (61),the airfoil body (61) 35 having external walls (68;268;268) extending between a leading edge (65) and a trailing edge (67;167;267); providing a baffle (72...272), the baffle (72...272) including a baffle body (74) defining an internal 40 passage (76), sidewalls (78) of the baffle body (74) defining a first contour; defining a cavity (64,64A-C), the cavity (64,64A- C) extending inwardly from the external walls (68;168;268) to define a second contour com- 45 plementary to the first contour; and inserting the baffle (72...272) into the cavity (64,64A-C).

14. The method as recited in claim 13, wherein: 50

the step of defining the cavity (64,64A-C) in- cludes removing material from the trailing edge (67;167;267) to define an opening (81) to the cavity (64,64A-C); and 55 the sidewalls (78) of the baffle body (74) are spaced apart by an exit wall (86;186;286) to de- fine one or more exit ports (82;182;282) situated

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