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I I l l 1 1 l l 1 1 1 1 1 1 Ill I l l 1 1 I l l 1 1 I I I I II I I I I 1 1 1 1 1 1 1 1 1 1 1 1 1 1 1 11111 1 l l 1 1 1 I l l 1 1 1 1 1 I l l 1
US0066 19030B 1
(12) United States Patent (io) Patent No.: US 6,619,030 B1
Seda et al. (45) Date of Patent: Sep. 16,2003
4,965,994 A 1011990 Ciokajlo (54) AIRCRAFT ENGINE WITH INTER-TURBINE 4,976,102 A 1211990 Taylor ENGINE FRAME SUPPORTED COUNTER 5,160,251 A 1111992 Ciokajlo ROTATING LOW PRESSURE TURBINE 5,177,957 A 111993 Grieb ROTORS 5,307,622 A * 511994 Ciokajlo et al. ......... 60139.162 (75) Inventors: Jorge F. Seda, Cincinnati, OH (US); 5,361,580 A 1111994 Ciokajlo Lawrence W. Dunbar, Cincinnati, OH 5,409,184 A 411995 Udal1 et al.
(US); Philip R. Gliebe, Waynesville, 5,438,756 A * 811995 Halchak et al. ............ 291889.2 OH (US); Peter N. Szucs, West 5,443,590 A 811995 Ciokajlo Chester, OH (US); John C. Brauer, 5,452,575 A 911995 Freid 5,813,214 A 911998 Monk et al.
Johnson, Hamilton, OH (US); Thomas Moniz, Loveland, OH (US); Gregory T. Steinmetz, Cincinnati, OH (US) * cited by examiner Primary Examinerxheryl J. Tyler (73) Assignee: General Electric Company, Assistant Examiner-John F. Belena Schenectady, NY (US) (74) Attorney, Agent, or F i r m q a t h a n D. Herkamp; ( * ) Notice: Subject to any disclaimer, the term of this Steven J. Rosen patent is extended or adjusted under 35 (57) ABSTRACT U.S.C. 154(b) by 68 days.
An aircraft gas turbine engine assembly includes an inter- (21) Appl. No.: 10/087,681 turbine frame axially located between high and low pressure turbines. Low pressure turbine has counter rotating low (22) Filed: Mar. 1, 2002 pressure inner and outer rotors with low pressure inner and outer shafts which are at least in part rotatably disposed (51) Int. C1.7 .................................................. F02K 3/02 co-axially within a high pressure rotor. Inter-turbine frame (52) U.S. C1. ...................... 60/226.1; 60139.162; 601792 includes radially spaced apart radially outer first and inner (58) Field of Search ........................... 60139.162, 226.1, second structural rings disposed co-axially about a center- 601268, 792, 796, 797 line and connected by a plurality of circumferentially spaced apart struts. Forward and aft sump members having forward (56) References Cited and aft central bores are fixedly joined to axially spaced U.S. PATENT DOCUMENTS apart forward and aft portions of the inter-turbine frame.
Low pressure inner and outer rotors are rotatably supported 4,558,564 A 1211985 Bouiller et al.
by a second turbine frame bearing mounted in aft central 4,758,129 A 711988 Strock bore of aft sump member. Amount for connecting the engine 4,790,133 A 1211988 Stuart to an aircraft is located on first structural ring.
4,809,498 A 311989 Giffin, I11 et al.
4,860,537 A 811989 Taylor 4,951,461 A 811990 Butler 21 Claims, 8 Drawing Sheets
U S . Patent Sep. 16,2003 Sheet 1 of 8 US 6,619,030 B1
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U S . Patent Sep. 16,2003 Sheet 8 of 8 US 6,619,030 B1
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SUMMARY OF THE INVENTION AIRCRAFT ENGINE WITH INTER-TURBINE ENGINE FRAME SUPPORTED COUNTER An aircraft gas turbine engine assembly includes a high ROTATING LOW PRESSURE TURBINE pressure rotor including a high pressure turbine, a low pressure turbine having counter rotating low pressure inner ROTORS S and outer rotors located aft of the high pressure rotor, and an inter-turbine frame axially located between the high and low This invention herein described was made in the perfor- pressure turbines. The low pressure inner and outer rotors mance of work under a NASA contract and is subject to the including low pressure inner and outer shafts which are at provisions of section 305 of the National Aeronautics and least in part rotatably disposed co-axially with and radially Space Act of 1958, Public Law 85-568 (72 Stat. 435; 42 inwardly of the high pressure rotor. The inter-turbine frame USC 2457).
has a first structural ring, a second structural ring disposed BACKGROUND OF THE INVENTION co-axially with and radially spaced inwardly of the first structural ring about a centerline, and a plurality of circum- 1. Field of the Invention ferentially spaced apart struts extending radially between the The invention relates to aircraft gas turbine engines with first and second structural rings. Forward and aft sump 1s counter rotating low pressure turbine rotors and, particularly, members, having forward and aft central bores respectively, for such engines having inter-turbine frames that support the are fixedly joined to axially spaced apart forward and aft counter rotating low pressure turbine rotors in bearings and portions of the inter-turbine frame by forward and aft are used to mount the engines to the aircraft.
bearing support structures, respectively. The low pressure 2. Description of Related Art inner and outer rotors are rotatably supported by an aftward- A gas turbine engine of the turbofan type generally most low pressure rotor support bearing mounted in the aft includes a forward fan and booster compressor, a middle central bore of the aft sump member. The high pressure rotor core engine, and an aft low pressure power turbine. The core is aftwardly radially rotatably supported by a fifth bearing engine includes a high pressure compressor, a combustor mounted in the forward bearing support structure. A frame and a high pressure turbine in a serial flow relationship. The connecting means for connecting the engine to an aircraft is high pressure compressor and high pressure turbine of the 2s located on the first structural ring. In an exemplary embodi- core engine are interconnected by a high pressure shaft. The ment of the invention, the connecting means includes at least high pressure compressor, turbine, and shaft essentially form one U-shaped clevis.
the high pressure rotor. The high pressure compressor is An outlet guide vane assembly supports a row of outlet rotatably driven to compress air entering the core engine to guide vanes that extend radially between a low pressure a relatively high pressure. This high pressure air is then turbine casing structurally connected to the inter-turbine mixed with fuel in the combustor and ignited to form a high frame and an annular box structure. A cover plate is bolted energy gas stream. The gas stream flows aft and passes to the annular box structure. A rotatable annular outer drum through the high pressure turbine, rotatably driving it and the rotor is drivingly connected to a first fan blade row and a first high pressure shaft which, in turn, rotatably drives the booster by the low pressure inner shaft. A rotatable annular compressor.
inner drum rotor is drivingly connected to a second fan blade 3s The gas stream leaving the high pressure turbine is row and a second booster by the low pressure outer shaft, the expanded through a second or low pressure turbine. The low first and second boosters are axially located between the first pressure turbine rotatably drives the fan and booster com- and second fan blade rows.
pressor via a low pressure shaft, all of which form the low A bypass duct radially bounded by a fan casing and an pressure rotor. The low pressure shaft extends through the annular radially inner bypass duct wall surrounds the first high pressure rotor. Some low pressure turbines have been and second boosters and a radially outer portion of the designed with counter rotating turbines that power counter second fan blade row is radially disposed within the bypass rotating fans and booster or low pressure compressors. U.S.
duct. The engine assembly has a fan inlet hub to tip radius Pat. Nos. 4,860,537, 5,307,622 and 4,790,133 disclose ratio in a range between 0.20 and 0.35, a bypass ratio in a counter rotating turbines that power counter rotating fans range of 5-15, an operational fan pressure ratio in a range of and booster or low pressure compressors. Most of the thrust 4s 1.4-2.5, and a sum of operational fan tip speeds of the first produced is generated by the fan. Engine frames including and second fan blade rows in a range of 1000 to 2500 feet fan and turbine frames are used to support and carry the per second. The high pressure compressor is designed and bearings which, in turn, rotatably support the rotors. Bearing operable to produce a compressor pressure ratio in a range support frames are heavy and add weight, length, and cost to of about 15-30 and overall pressure ratio in a range of about the engine.
4&65.
so Large modern commercial turbofan engines have higher Further embodiments of the invention include a second operating efficiencies with higher b y p a s s ratio seal in sealing arrangement between forward ends of the low configurations, larger transition ducts between low pressure pressure turbine casing and the outer drum rotor, a third seal and high pressure turbines. The frames, especially those in sealing arrangement between the low pressure turbine located in the engine hot section, are complex and expen- casing and a final stage of the low pressure turbine blade 5s sive. Other mid-size turbofan engines eliminate one frame rows which is bolted to an aft end of the outer drum rotor, by providing HP rotor support through a differential bearing and a first seal in sealing arrangement between the second arrangement in which the high pressure rotor rides on the fan and the fan frame. The seals are brush seals, however in low pressure rotor with an inter-shaft or differential bearing other embodiments the seals may be non-contacting seals or between them. New commercial engine designs are incor- a combination of brush seals and non-contacting seals. The porating counter rotating rotors for improved turbine effi- non-contacting seals may be aspirating seals or face seals.
ciency. Counter rotating rotors can have a detrimental BRIEF DESCRIPTION OF THE DRAWINGS impact on high pressure ratio component clearances, espe- The foregoing aspects and other features of the invention cially in the hot section which rely on tight clearance control are explained in the following description, taken in connec- to provide fuel efficiency benefits. Consequently, a need 65 tion with the accompanying drawings where: exists for engine and bearing support that will reduce engine length, weight, and cost and still provide low tip clearance FIG. 1 is a longitudinal sectional view illustration of losses. exemplary embodiment of an aircraft turbofan gas turbine
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engine with a counter rotating low pressure turbine sup- rounds the counter rotating first and second boosters 16 and ported by an inter-turbine frame axially located between the 17 and an inlet duct 19 to the high pressure compressor 18 of the core engine 25. The bypass duct 21 is radially low pressure turbine and a high pressure turbine.
bounded by a fan casing 11 and an annular radially inner FIG. 2 is a longitudinal sectional view illustration of an bypass duct wall 9. The radially inner bypass duct wall 9 alternative exemplary embodiment of an aircraft turbofan includes a rotatable wall section 22 fixedly mounted to the gas turbine engine with a two stage high pressure turbine and second fan blade row 15 and from which the second booster a counter rotating low pressure turbine supported by an 17 depends radially inwardly. A radially outer portion 23 of inter-turbine frame axially located between the low pressure the second fan blade row 15 is radially disposed within the turbine and a high pressure turbine.
bypass duct 21.
FIG. 3 is an enlarged view illustration of the inter-turbine Referring now to FIGS. 1 and 3, the counter rotating low frame and counter rotating low pressure turbine rotors of the pressure turbine 26 includes an annular outer drum rotor 136 engine in FIG. 1. rotatably mounted to a low pressure inner shaft 130 by an aft low pressure inner conical shaft extension 132. The outer FIG. 4 is an enlarged view illustration of a fan frame and 136 includes a plurality of low pressure turbine drum rotor forward bearings and sump of the engine in FIG. 1.
15 blade rows 138 extending radially inwardly therefrom and FIG. 5 is an enlarged view illustration of a radially inner 136 is axially spaced from each other. The drum rotor forward brush seal in the counter rotating low pressure cantilevered off of a final stage 139 of the low pressure turbine in FIG. 1.
turbine blade rows 138 which is bolted to the aft low FIG. 6 is an enlarged view illustration of a radially outer pressure inner conical shaft extension 132. The counter forward brush seal in the counter rotating low pressure rotating low pressure turbine 26 also includes an annular low turbine in FIG. 1.
20 pressure inner drum rotor 146 rotatably mounted to a low pressure outer shaft 140 by an aft low pressure outer conical FIG. 7 is an enlarged view illustration of an aft brush seal shaft extension 142. The inner drum rotor 146 includes a in the counter rotating low pressure turbine in FIG. 1.
plurality of second low pressure turbine blade rows 148 FIG. 8 is an enlarged view illustration of a radially inner extending radially outwardly therefrom and axially spaced forward non-contacting seal in the counter rotating low 2s from each other. The first low pressure turbine blade rows pressure turbine in FIG. 1.
138 are interdigitated with the second low pressure turbine FIG. 9 is an enlarged view illustration of a radially outer blade rows 148.
forward non-contacting seal in the counter rotating low The low pressure outer shaft 140 drivingly connects the pressure turbine in FIG. 1.
inner drum rotor 146 to the second fan blade row 15 and the FIG. 10 is an enlarged view illustration of an aft non- 3o second booster 17. The second fan blade row 15 is connected contacting seal in the counter rotating low pressure turbine to the low pressure outer shaft 140 by a forward conical in FIG. 1.
outer shaft extension 143. The low pressure outer shaft 140, the inner drum rotor 146, the second fan blade row 15, and DETAILED DESCRIPTION OF THE the second booster 17 are major components of a low INVENTION pressure outer rotor 202. The low pressure inner shaft 130 Illustrated schematically in FIG. 1 is a first exemplary 35 drivingly connects the outer drum rotor 136 to the first fan turbofan gas turbine engine 10 circumscribed about an blade row 13 and the first booster 16. The first fan blade row engine centerline 8 and having a fan section 12 which 13 is connected to the low pressure inner shaft 130 by a receives inlet airflow of ambient air 14. The fan section 12 forward conical inner shaft extension 133. The low pressure has counter rotating first and second fans 4 and 6 including inner shaft 130, the outer drum rotor 136, the first fan blade first and second fan blade rows 13 and 15 and counter 40 row 13, and the first booster 16 are major components of a rotating first and second boosters 16 and 17, respectively. low pressure inner rotor 200.
The counter rotating first and second boosters 16 and 17 are The first booster 16 includes an annular first booster rotor axially located between the counter rotating first and second section 166 including the rotatable wall section 22 from fan blade rows 13 and 15, an arrangement which provides which axially spaced apart first booster blade rows 168 reduced noise emanating from the fan section 12. Following 45 extend radially inwardly. The annular first booster rotor the fan section 12 is a high pressure compressor (HPC) 18, section 166 is illustrated as being integrally bladed in a a combustor 20 which mixes fuel with the air 14 pressurized manner similar to an integrally bladed disk, commonly by the HPC 18 for generating combustion gases which flow referred to as a Blisk, or an integrally bladed rotor which has downstream through a high pressure turbine (HPT) 24, and been used in conventional rotors because they are light- a counter rotating low pressure turbine (LPT) 26 from which so weight and allow no blade attachment leakage. The operat- the combustion gases are discharged from the engine 10. The ing low speeds of the boosters and the low weight integrally engine 10 is designed such that the last stage of the second bladed disk design of the first booster rotor section 166 helps booster 17 and, in the exemplary embodiment, the second minimize stresses and deflections of the first booster rotor fan blade row 15 are counter rotatable with respect to the section 166.
high pressure compressor 18. This reduces the sensitivity of The second booster 17 includes an annular second booster the engine 10 to airflow inlet distortion of the fan section 12.
rotor section 170 from which axially spaced apart second It also reduces mutual sensitivity to rotating stall cells in the booster blade rows 172 extend radially outwardly. A radially other rotors.
inner portion 28 of the second fan blade row 15 is radially A high pressure shaft 27 joins the HPT 24 to the HPC 18 disposed within the inlet duct 19 and rotates with the second to substantially form a first or high pressure rotor 33. The booster 17 and therefore is considered part of the second high pressure compressor 18, combustor 20, and high pres- 6o booster 17 and a second booster blade row 172. The first and sure turbine 24 collectively are referred to as a core engine second booster blade rows 168 and 172 are interdigitated 25 which includes, for the purposes of this patent, the high and are counter rotating. The first and second fan blade rows pressure shaft 27. The core engine 25 is modular such that 13 and 15 are fixedly attached to the first and second booster as a single unit it can be independently replaced separate rotor sections 166 and 170, respectively. The low pressure from the other parts of the gas turbine.
65 inner and outer shafts 130 and 140, respectively, are at least, Abypass duct 21 radially, bounded by a fan casing 11 and in part, rotatably disposed co-axially with and radially a rotatable annular radially inner bypass duct wall 9, sur- inwardly of the high pressure rotor 33.
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The engine 10 has frame structure 32 including a forward 140, is rotatably supported radially by a third bearing 76 or fan frame 34 connected by an engine casing 45 to a within the inter-turbine frame 60. The third bearing 76 is mid-engine or inter-turbine frame 60. The second fan blade disposed between an aft bearing support structure 97 row is axially located close to struts 35 of the fan frame 34 attached to an aft portion 110 of the inter-turbine frame 60 5 and a forward inner extension 190 of the aft low pressure and so the leading edges of struts 35 are swept or leaned axially aftwardly to reduce noise. The engine 10 is mounted outer conical shaft extension 142. The low pressure outer within or to an aircraft such as by a pylon (not illustrated) rotor 202 is most aftwardly rotatably supported by the third which extends downwardly from an aircraft wing. The bearing 76 which is thus referred to as an aftwardmost low inter-turbine frame 60 includes a first structural ring 86, pressure rotor support bearing. The inter-turbine frame 60 of is axially located between the HPT 24 which may be a casing, disposed co-axially about the the present invention centerline 8. The inter-turbine frame 60 further includes a lo and the LPT26 and thus substantially supports the entire low second structural ring 88 disposed co-axially with and pressure turbine 26.
radially spaced inwardly of the first structural ring 86 about The low pressure inner rotor 200, by way of the aft low the centerline 8. The second structural ring 88 may also be pressure inner conical shaft extension 132 connected to the referred to as a hub. A plurality of circumferentially spaced low pressure inner shaft 130, is rotatably supported radially apart struts 90 extend radially between the first and second by the aft low pressure outer conical shaft extension 142 of rings 86 and 88 and are fixedly joined thereto. The struts 90 the low pressure outer rotor 202. A differential bearing 144 are hollow in the exemplary embodiment of the invention (also referred to as an inter-shaft bearing) is disposed illustrated herein but, in other embodiments, the struts may between an aft inner extension 192 of the aft low pressure not be hollow. Because the inter-turbine frame 60 is axially outer conical shaft extension 142 and an outer extension 194 located between the HPT 24 and the LPT 26 of the high 20 of the aft low pressure inner conical shaft extension 132.
pressure rotor 33 and the low pressure inner and outer rotors This allows the low pressure inner and outer rotors 200 and 200 and 202, it is referred to as an inter-turbine frame also 202 to counter rotate.
sometimes referred to as a mid-engine frame. h. inter- Referring to FIG. 1, a forward high pressure end 70 of the turbine transition duct 114 between the HPT 24 and the LPT high pressure compressor 18 of the high pressure rotor 33 is 26 passes through the inter-turbine frame 60. 2s radially rotatably supported by a bearing assembly 80 The engine is mounted to the aircraft at a forwardly mounted in a bearing assembly support structure 82 attached located fan frame forward mount 118 on the fan frame 34 to the fan frame 34. Referring more particularly to FIG. 3, and at an aftwardly located turbine frame aft mount 120 on an aft end 92 of the high pressure rotor 33 is aftwardly the inter-turbine frame 60. The engine 10 may be mounted radially rotatably supported by a fifth bearing 94 mounted in below an aircraft wing by a pylon at the forward mount 118 3o a forward bearing support structure 96 attached to a forward and the aft mount 120 spaced axially downstream from the portion 108 of the inter-turbine frame 60. The forward and forward mount 118. The aft mount 120 is used to fixedly join aft bearing support structures 96 and 97 which are fixedly the inter-turbine frame 60 to a platform which is fixedly joined or attached to the forward and aft portions 108 and joined to the pylon. In the exemplary embodiment of the 110, respectively, of the inter-turbine frame 60 and thus are invention illustrated herein, the aft mount 120 includes a spaced axially apart. The forward and aft portions 108 and U-shaped clevis 122. Conventional mounts often use a set of 3s 110, respectively, of the inter-turbine frame 60 are separated circumferentially spaced apart U-shaped devises 122 (only by the second structural ring 88.
one of the U-shaped devises is shown in the cross-sectional Forward and aft sump members 104 and 106 are joined to illustrations in the FIGS.) on the inter-turbine frame 60. The the inter-turbine frame 60 and carried by forward and aft U-shaped devises 122 are designed to be connected by a set bearing support structures 96 and 97. The forward and aft of pins to a set of links. The links are connected to a platform 40 sump members 104 and 106 support the fifth bearing 94 and on the bottom of the pylon. The U-shaped devises 122 are the third bearing 76 in forward and aft cylindrical central one type of frame connecting means for connecting the bores 84 and 85, respectively, of the sump members. The engine to an aircraft. Other types of mounting means besides fifth bearing 94 and the third bearing 76 have forward and devises are known in the aircraft industry and can be utilized aft fixed outer races 176 and 178 that are fixedly connected to mount the frame of the present invention and the engine 4s to the forward and aft bearing support structures 96 and 97, to the aircraft. respectively.
Referring more particularly to FIG. 4, the low pressure Located aft of the LPT 26 is an outlet guide vane outer rotor 202, by way of the forward conical outer shaft assembly 150 which supports a stationary row of outlet extension 143, is rotatably supported axially and radially guide vanes 152 that extend radially inwardly between a low from the fan frame 34 by an aft thrust bearing 43 mounted so pressure turbine casing 54 and an annular box structure 154.
in a first bearing support structure 44 and a second bearing The outlet guide vane assembly 150 deswirls gas flow 36, a roller bearing, mounted in a second bearing support exiting the LPT 26. The low pressure turbine casing 54 structure 47. The low pressure inner rotor 200, by way of the connected is bolted to the engine casing 45 at the end of the forward conical inner shaft extension 133, is rotatably inter-turbine transition duct 114 between the HPT24 and the supported axially and radially from the fan frame 34 by a LPT 26. A dome-shaped cover plate 156 is bolted to the 5s forward differential thrust bearing 55 which is mounted annular box structure 154. The outlet guide vane assembly between a forwardly extending extension 56 of the forward 150 is not referred to and does not function as a frame conical outer shaft extension 143 and the forward conical because it does not rotatably support any of the engine’s inner shaft extension 133. The low pressure inner rotor 200 rotors.
is further rotatably supported radially from the fan frame 34 The aft sump member 106 has a first radius R 1 from the by a forward differential bearing 208, a roller bearing, 6o engine centerline 8 that is substantially greater than a second between the low Pressure inner shaft 130 and the low radius R2 of the forward sump members 104. The first radius Pressure outer shaft 140. The first and second bearing R 1 may be in a range of 150 to 250 percent larger than the support structures 44 and 47 are fixedly attached to the fan second radius R2. The aft sump member 106 is located frame 34.
radially from the engine centerline 8 a distance that is Referring more particularly to FIG. 3, the low pressure 65 substantially greater than the distance in similarly sized prior outer rotor 202, by way of the aft low pressure outer conical engines. This helps stiffen the third bearing 76 in the aft shaft extension 142 connected to the low pressure outer shaft central bore 85 as does the forward and aft bearing support
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structures 96 and 97 being axially spaced apart and fixedly allows a smaller fan diameter when compared to a larger joined to forward and aft portions 108 and 110 of the ratio. However, fan inlet hub to tip radius ratio levels are inter-turbine frame 60, respectively, and separated by the constrained by the ability to design a disk to support the second structural ring 88. These design features improve rotating fan blades. The fan blades in the exemplary embodi- 5 ment illustrated herein are made of lightweight composite maneuver clearances by increasing stiffness of the support of the third bearing 76 and the differential bearing 144 which materials or aluminum and rotor fan tip speeds are designed rotatably support the inner drum rotor 146 and the annular so that a fan disk 240 can be designed for the fan inlet hub outer drum rotor 136, respectively, of the LPT 26. to tip radius ratio to be as low as 0.20. The low fan inlet hub to tip radius ratio allows low slopes and short lengths of the Illustrated schematically in FIG, 2 is an alternative tur- bofan gas turbine engine 10 configuration having a two stage core engine transition duct 242 between the fan section 12 high pressure turbine 324 having two stages of high pressure and the high Pressure ComPressor 18 and ofthe inter-turbine turbine blades 326 and a row of high pressure turbine vanes transition duct 114 between the HPT 24 and the LPT 26.
328 as opposed to the turbofan gas turbine engine 10 Oil lubrication and scavenging of oil for the differential illustrated in FIGS. 1 and 3 in which the HPT 24 is a single bearing 144 is accomplished by routing oil through supply stage high pressure turbine with a single stage of HPT blades and return orifices 220 and 222, respectively, at low stress 326. This illustrates how significant engine thrust growth location on the aft low pressure outer conical shaft extension can be achieved with no change in fan diameter D. Thrust 142 as illustrated in FIGS. 1, 3, and 4. This feature helps grow may be achieved by increasing a design fan pressure strengthen the rotors and allow the use of only two bearing ratio across the second fan blade rows 15 stage with the support frames, the fan and inter-turbine frames, for the 3 same fan casing or fan diameter D as with the single stage spool design with counter rotating low pressure turbines and high pressure turbine. This allows for a family of engines 20 rotors.
based on substantially the same engine architecture and sealing of the counter rotating low pressure rotors to basic components. The fan diameter D of all the engines 10 booster and LPT cases is enhanced by the use of brush seals in the family would be substantially the same. At least two or other high seals, Illustrated in FIG, 5 is a first different engine models in the family would have either the brush seal 223 disposed in sealing arrangement between the one stage high Pressure turbine 24 O r the two stage high 2s second fan 6 and the fan frame 34. Illustrated in FIG. 6 is a pressure turbine 324 of the core engine 25.
second brush seal 224 disposed in sealing arrangement The high pressure compressor 18 of turbofan gas turbine between forward ends 226 of the low pressure turbine casing engine 10 of the present invention is operable and designed 54 and the outer drum rotor 136. Illustrated in FIG. 7 is a to operate with a relatively high compressor pressure ratio in third brush seal 225 disposed in sealing arrangement a range of about 15-30 and an overall pressure ratio in a 30 between an aft end 228 of the low pressure turbine casing 54 range of about 4e6.5. The compressor pressure ratio is a and the final stage 139 of the low pressure turbine blade measure in the rise of pressure across just the high pressure rows 138 which is bolted to the outer drum rotor 136. An compressor 18. The overall pressure ratio is a measure in the alternative to the brush seals are non-contacting seals such rise of pressure across the fan all the way through the high as aspirating seals or face seals in some or all of the above pressure compressor 18, i.e., it is a ratio of pressure exiting 3s locations. Illustrated in FIG. 8 is a first non-contacting seal the high pressure compressor divided by pressure of ambient 244 disposed in sealing arrangement between the second fan air 14 entering the fan section 12. The high pressure com- 6 and the fan frame 34. Illustrated in FIG. 9 is a second pressor 18 is illustrated having six high pressure stages 48 non-contacting seal 246 disposed in sealing arrangement and three variable vane stages 50 for the first four of the high between forward end 226 of the low pressure turbine casing pressure stages 48. Less than four variable vane stages 50 54 and the outer drum rotor 136. Illustrated in FIG. 10 is a may be used. The high pressure compressor 18 has a 40 third non-contacting seal 248 disposed in sealing arrange- relatively small number of the high pressure stages 48 and ment between an aft end 228 of the low pressure turbine the invention contemplates using between 6-8 of the high casing 54 and the final stage 139 of the low pressure turbine pressure stages and about four of the variable vane stages 50 blade rows 138 which is bolted to the outer drum rotor 136.
or less. This makes for a short engine while still having a In other embodiments the seals may be a combination of high overall pressure ratio in a range of 40-65.
45 brush seals and non-contacting seals.
The engine has a design bypass ratio in a range of 5-15 The present invention has been described in an illustrative and a design fan pressure ratio in a range of 1.4-2.5. The manner. It is to be understood that the terminology which counter rotating first and second fan blade rows 13 and 15 has been used is intended to be in the nature of words of are designed to operate with tip speeds that, for the two blade description rather than of limitation. While there have been rows, sum to a range of about 1000 to 2500 ftisec which so described herein, what are considered to be preferred and allows the use of light weight composite fan blades. Light exemplary embodiments of the present invention, other weight, uncooled, high temperature capability, counter rotat- modifications of the invention shall be apparent to those ing ceramic matrix composite (CMC) airfoils may be used skilled in the art from the teachings herein and, it is, in the counter rotating low pressure turbine 26. Thus the therefore, desired to be secured in the appended claims all engine 10 and the fan section 12 may be described as having ss such modifications as fall within the true spirit and scope of a sum of operational fan tip speeds of the first and second fan the invention.
rows l3 and l5 in a range Of loo0 to 2500 feet per Accordingly, what is desired to be secured by Letters second.
Patent of the United States is the invention as defined and Illustrated in FIG. 1, is a tip radius RT, as measured from differentiated in the following claims: the engine centerline 8 to a fan blade tip 230 of the first fan What is claimed is: 6o blade row 13 and a hub radius RH as measured from the 1, ~n aircraft gas turbine engine assembly comprising: engine centerline 8 to a rotor hub 234 of the low pressure a high pressure rotor including a high pressure turbine, inner rotor 200 at an entrance 235 to the inlet duct 19 to the a low pressure turbine having counter rotating low pres- high pressure compressor 18 of the core engine 25, The sure inner and outer rotors located aft of said high engine 10 of the present invention may be designed with a Pressure small fan inlet hub to tip radius ratio (RHIRT) in a range 65 an inter-turbine frame axially located between said high between 0.20 and 0.35. For a given set of fan inlet and inlet duct annulus areas a low fan inlet hub to tip radius ratio and low pressure turbines,
US 6,619,030 B3
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said low pressure inner and outer rotors including low 9. An assembly as claimed in claim 6 further comprising: pressure inner and outer shafts which are at least in part a fan inlet hub to tip radius ratio in a range between 0.20 rotatably disposed co-axially with and radially and 0.35, inwardly of said high pressure rotor, a high pressure compressor of said high pressure rotor S said inter-turbine frame comprising; drivenly connected to said high pressure turbine by a a first structural ring, high pressure shaft, a second structural ring disposed co-axially with and said high pressure compressor having an overall operating radially spaced inwardly of said first structural ring pressure ratio in a range-of about 40-65, and about a centerline, a plurality of circumferentially spaced apart struts a bypass ratio in a range of 5-15, an operational fan extending radially between said first and second pressure ratio in a range of 1.4-2.5, and a sum of structural rings, operational fan tip speeds of said first and second fan forward and aft sump members having forward and aft blade rows in a range of 1000 to 2500 feet per second.
central bores, 10. An assembly as claimed in claim 9 wherein said frame 1s said forward and aft sump members fixedly joined to connecting means includes at least one U-shaped clevis.
axially spaced apart forward and aft portions of said 11. An assembly as claimed in claim 10 further compris- inter-turbine frame by forward and aft bearing sup- ing a cover plate bolted to said annular box structure.
port structures respectively, 12. An assembly as claimed in claim 6 further comprising said low pressure inner and outer rotors rotatably a second seal in sealing arrangement between forward ends supported by an aftwardmost low pressure rotor of the low pressure turbine casing and the outer drum rotor.
support bearing mounted in said aft central bore of 13. An assembly as claimed in claim 12 further compris- said aft sump member, ing a third seal in sealing arrangement between said low said high pressure rotor is aftwardly radially rotatably pressure turbine casing and a final stage of said low pressure supported by a fifth bearing mounted in said forward turbine blade rows which is bolted to an aft end of said outer 2s bearing support structure, and drum rotor.
a frame connecting means for connecting said engine to 14. An assembly as claimed in claim 13 further compris- an aircraft located on said first structural ring.
ing a first seal in sealing arrangement between said second 2. An assembly as claimed in claim 1 wherein said frame fan and said fan frame.
connecting means includes at least one U-shaped clevis.
15. An assembly as claimed in claim 14 wherein said seals 3. An assembly as claimed in claim 1 further comprising are brush seals.
an outlet guide vane assembly supporting a row of outlet 16. An assembly as claimed in claim 14 wherein said seals guide vanes that extend radially between a low pressure are non-contacting seals.
turbine casing structurally connected to said inter-turbine 17. An assembly as claimed in claim 14 wherein said seals frame and an annular box structure.
are brush seals or non-contacting seals.
4. An assembly as claimed in claim 3 wherein said frame 3s 18. An assembly as claimed in claim 14 wherein said connecting means includes at least one U-shaped clevis.
non-contacting seals are aspirating seals or face seals.
5. An assembly as claimed in claim 4 further comprising 19. An assembly as claimed in claim 14 further compris- a cover plate bolted to said annular box structure.
ing: 6. An assembly as claimed in claim 3 further comprising: a fan inlet hub to tip radius ratio in a range between 0.20 a rotatable annular outer drum rotor drivingly connected and 0.35, to a first fan blade row and a first booster by said low a high pressure compressor of said high pressure rotor pressure inner shaft, drivenly connected to said high pressure turbine by a a rotatable annular inner drum rotor drivingly connected high pressure shaft, to a second fan blade row and a second booster by said said high pressure compressor having an overall operating low pressure outer shaft, 4s pressure ratio in a range of about 40-65, and said first and second boosters are axially located between a bypass ratio in a range of 5-15, an operational fan said first and second fan blade rows, pressure ratio in a range of 1.4-2.5, and a sum of a bypass duct radially bounded by a fan casing and an operational fan tip speeds of said first and second fan annular radially inner bypass duct wall surrounding blade rows in a range of 1000 to 2500 feet per second.
so said first and second boosters, and 20. An assembly as claimed in claim 19 wherein said a radially outer portion of said second fan blade row is frame connecting means includes at least one U-shaped radially disposed within said bypass duct. clevis.
21. An assembly as claimed in claim 20 further compris- 7. An assembly as claimed in claim 6 wherein said frame connecting means includes at least one U-shaped clevis. ing a cover plate bolted to said annular box structure.
5s 8. An assembly as claimed in claim 7 further comprising
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a cover plate bolted to said annular box structure.