Patent Application
20180142564
The present invention relates generally to gas turbine engine turbine nozzles and shrouds.
In a typical gas turbine engine, air is compressed in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. The gases flow downstream through a high pressure turbine (HPT) having one or more stages including one or more HPT turbine nozzles, shrouds, and rows of HPT rotor blades. The gases then flow to a low pressure turbine (LPT) having one or more stages of respective LPT turbine nozzles, shrouds, and LPT rotor blades. The HPT and LPT turbine nozzles include a plurality of circumferentially spaced apart stationary nozzle vanes extending radially between outer and inner bands. Each nozzle vane may have a hollow airfoil through which cooling air is flowed. Cooling air for each vane can be fed through a single spoolie located radially outwardly of the outer band of the nozzle. In some vanes subjected to higher temperatures, such as the HPT vanes for example, an impingement baffle may be inserted in each hollow airfoil to supply cooling air to the airfoil.
The turbine rotor stage includes a plurality of circumferentially spaced apart rotor blades extending radially outwardly from a rotor disk. Turbine nozzles are located axially forward of a turbine rotor stage. The turbine shrouds are located radially outward from the tips of the turbine rotor blades so as to form a radial clearance between the rotor blades and the shrouds. The shrouds are held in position by shroud hangers which are supported by flanges engaging with annular casing flanges.
The turbine nozzles, shrouds, and shroud hangers are typically formed in arcuate segments. Each nozzle segment typically has two or more vanes joined between an outer band segment and an inner band segment. Each nozzle segment and shroud hanger segment is typically supported at its radially outer end by flanges attached to an annular outer and/or inner casing. Each vane has a cooled airfoil disposed between radially inner and outer band panels which form the inner and outer bands. In some designs, the airfoil, inner and outer band portions, flange portion, and intake duct are cast together such that the vane is a single casting. In some other designs, the vane airfoils are inserted in corresponding openings in the outer band and the inner band and brazed along interfaces to form the nozzle segment.
A leakage path between the nozzle and shroud is a source of engine fuel efficiency and performance losses. The leakage path at the outer band can cause a hot hook on the case which can be a life limiting low cycle fatigue feature. During engine operation, the nozzle radially chords and it is desirable to reduce or prevent this chording. Chording of the bands can lead to interference with angel wing overlap. Angel wing elements provide sealing between hot turbine flow stream and cool cavity within the rotor. Inter-turbine seals on the engine's turbine transition duct are subject to axial excursions which should be minimized throughout all operating conditions. Chording occurs on the bands due to the high temperature at the band flowpath combating the colder temperatures on the non-flowpath sides of the bands, particularly, the flanges. Chording of the bands is bowing away from the flowpath.
Certain turbines have a cantilevered nozzle mounted and cantilevered from the outer band. Some nozzle segments are configured with multiple airfoils. Two vane designs may be referred to as doublets and three vane designs may be referred to as triplets. Doublets and triplets offer performance advantages in reducing split-line leakage flow between vane segments. However, the longer chord length of the bands and mounting structure compromises the durability of the multiple vane nozzle segments. The longer chord length causes an increase in chording and its adverse effects discussed above due to the higher displacement of the longer chord length activated by the radial thermal gradient through the band.
It is desirable to eliminate turbine leakage paths between the nozzle and shroud in order to improve engine fuel efficiency and performance. It is also desirable to reduce chording of bands or flanges and its negative effect or interference with angel wing overlap and sealing. It is desirable to reduce part count, decrease leakage, and enable the turbine case to be a simpler and lower cost configuration.
A gas turbine engine turbine nozzle assembly includes airfoils extending between annular coaxial inner and outer band segments circumscribed about a longitudinal or axial centerline axis, a shroud segment extending aft and downstream from the outer band segment, and the outer band segment and shroud segment being integral, monolithic, and integrally formed. One or more deflection limiters extending radially outwardly from the outer band segment limit amount of axial excursion at the inner band segments.
The outer band segment may extend axially between arcuate forward and middle outer flanges. The outer band segment and the shroud segment may be integrally formed by equiax, directionally solidified, or single crystal casting.
The assembly may include bonded together singlets. Each of the singlets includes one of the airfoils extending radially between arcuate inner and outer band singlet segments, the outer band singlet segment extending axially between arcuate forward and middle outer singlet flanges, the forward and middle outer singlet flanges extending radially outwardly from the outer band singlet segment, and the singlet including a shroud singlet segment extending aft and downstream from the outer band singlet segment to an aft singlet flange. The singlets may be integrally formed by equiax, directionally solidified, or single crystal casting.
The assembly includes one or more deflection limiters on the middle outer flange for limiting amount of axial excursion at the inner band segment. The one or more deflection limiters may include one or more outward projections respectively extending upwardly or radially outwardly from the middle outer flange.
In the embodiment with singlets, each of the outward projections may include two circumferentially adjacent outward complimentary portions on two circumferentially adjacent ones of the singlets and the singlets may be integrally formed by equiax, directionally solidified, or single crystal casting.
A gas turbine engine may include the assembly surrounded and supported by an annular outer casing. Inward projections may extend inwardly or radially inwardly from the casing at or about axial and circumferential locations of the deflection limiters.
The gas turbine engine may include a radially inner discourager or angel wing seal assembly sealingly disposed between an inner turbine duct and a high pressure turbine disk or rotor of the high pressure turbine upon which high pressure turbine blades are mounted. The inner turbine duct is attached to the assembly. The discourager or angel wing seal assembly includes an annular seal land extending forwardly or upstream from the inner turbine duct and is radially disposed between and in sealing relationship with aftwardly or downstream extending radially outer and inner angel wings attached to the high pressure turbine disk or rotor. The inner angel wing may be on an annular cooling plate attached to the high pressure turbine disk.
The subject matter which is regarded as the invention is particularly pointed out and distinctly claimed in the concluding part of the specification. The invention, in accordance with preferred and exemplary embodiments, together with further objects and advantages thereof, is described in the following detailed description taken in conjunction with the accompanying drawings in which:
Illustrated schematically in
Referring to
Turbine stator components such as low pressure turbine nozzles and shrouds are often manufactured in arcuate segments and then assembled together in the engine 10 forming the turbine components. Various joints or gaps are provided between annular assemblies of arcuate segments which must be suitably sealed for preventing leakage of the high pressure cooling air 29 into the turbine flowpath 27. An annular turbine nozzle assembly 32 includes a circular row of circumferentially adjacent turbine nozzle segments 40 having outer bands 35 integral, monolithic, and integrally formed with shroud segments 48 that eliminates these joints or gaps as further illustrated in
An exemplary embodiment of such a turbine nozzle segment 40 is further illustrated for the first low pressure turbine nozzle 24 in
Referring to
Each of the turbine nozzle assemblies 32 further includes a shroud segment 40 extending aft and downstream from a radially outer end 49 of the middle outer flange 72 to an aft flange 74. The exemplary embodiments of the outer band segment 38 and the shroud segment 48 illustrated herein are conical having substantially equal cone angles 50 and are spaced radially apart. The shroud segment 48 is spaced radially outward of the outer band segment 38. The shroud 220 is made from a circular array of the shroud segments 48. An annular shroud seal land 110 (honeycomb) extends radially inwardly of the shroud 220 and is made from a circular array of seal land segments 112 bonded to the shroud segments 48. Knife edge blade seals 114 extending radially outwardly from the low pressure turbine blades 23 sealingly engage the annular shroud seal land 110 as illustrated in
Referring to
Each turbine nozzle assembly 32 may be made from subassemblies 51 such as singlets 31 as illustrated herein in the exemplary embodiment of the assembly. Circumferentially adjacent singlets 31 in the turbine nozzle assembly 32 are joined together. The singlets 31 may be joined together by brazing, indicated by braze line 47. Each singlet 31 includes one of the airfoils 34 extending radially between arcuate inner and outer band singlet segments 237, 238. Note, that the subassemblies 51 may include two or more airfoils between the arcuate inner and outer band segments such as doublets having two airfoils or triplets having three airfoils. The outer band singlet segment 238 extends axially between arcuate forward and middle outer singlet flanges 270, 272 which extend radially outwardly from the outer band singlet segment 238. The singlet 31 further includes a shroud singlet segment 240 extending aft and downstream from a radially outer end 49 of the middle outer singlet flange 272 to an aft singlet flange 274. The exemplary embodiments of the outer band singlet segment 238 and the shroud singlet segment 240 illustrated herein are conical having substantially equal cone angles 50 and are spaced radially apart. The forward and middle outer singlet flanges 270, 272 and the aft singlet flange 274 extend circumferentially between circumferentially spaced apart first and second singlet edges 262, 264 of the singlet 31.
The cooling air 29 is at a higher pressure compared to that of the combustion gases 30 flowing through the low pressure turbine nozzle 24. The relatively cold cooling air 29 cools the outer band segment 38 and the forward and middle outer flanges 70, 72 attached thereto. The relatively hot combustion gases 30 in the turbine flowpath 27 produce a hot surface 56 along an inner side 58 of the outer band segment 38. The inner band segment 37 has a cold surface 52 along an inner side 58 of the inner band segment 37. The relatively hot combustion gases 30 in the turbine flowpath 27 produce a hot surface 56 along a radially outer side 54 of the inner band segment 37.
Chording of all the outer band segments 38 in the entire outer band 35 including the forward and middle outer flanges 70, 72 occurs due to a thermal gradient associated with the hot combustion gases 30. The hot combustion gases 30 impart high temperatures to the flowpath side of the outer band 35 including the outer band segments 38 and the forward and middle outer flanges 70, 72. The cooling air 29 imparts cold temperatures on the outer surfaces of the outer band including the outer band segment 38 and the forward and middle outer flanges 70, 72. The thermal gradient is also due to conduction between the outer casing 44 and the turbine nozzle assembly 32 which occurs at the forward and aft hooks 45, 46. The temperature gradient is exacerbated by the flanges radial height resulting in a higher thermal gradient.
The first low pressure turbine nozzles 24 experience high stresses at the interface between the airfoils 34 and the band segments, particularly at trailing edges TE of the airfoils 34. The high stress results in cracking at these locations. One of the highest contributors to this stress is the chording, particularly of the flanges. As the band undergoes chording (bowing away from the flowpath), the airfoils are pulled on, resulting in high stresses.
The turbine nozzle assembly 32 includes a deflection limiting means 60 for counteracting chording or flattening. The deflection limiting means 60 illustrated herein includes one or more deflection limiters 100 extending radially outwardly from the outer band 35 or more particularly from at least one of the outer band segments 38 for limiting amount of axial excursion at the inner band 36 or the inner band segments 37. One embodiment of the deflection limiting means 60 illustrated herein are deflection limiters 100 on each middle outer flange 72. Two deflection limiters 100 are on each middle outer flange 72 of the exemplary embodiment of the turbine nozzle assembly 32 illustrated herein. The deflection limiters 100 limit the amount of radial excursion at the outer band 35 which in turn limit the amount of axial excursion at the inner band 36 and the inner band segments 37. Referring to
A radially inner angel wing seal assembly 130 provides sealing between the inner turbine duct 128 and a high pressure turbine disk 148 or rotor of the high pressure turbine 19 upon which high pressure turbine blades 140 are mounted. The angel wing seal assembly 130 is a discourager seal and includes an annular seal land 142 extending forwardly or upstream from the inner turbine duct 128. The seal land 142 is radially disposed in sealing relationship with and between aftwardly or downstream extending radially outer and inner angel wings 144, 146 attached to the high pressure turbine disk 148 or rotor. The inner angel wing 146 may be on an annular cooling plate 150 attached to the high pressure turbine disk 148 or rotor upon which the high pressure turbine blades 140 are mounted. The deflection limiters 100 limit the amount of axial excursion at the inner band 36 and the inner band segments 37 and limits or prevents interference with between the inner turbine duct 128 and the high pressure turbine disk 148 or rotor and the high pressure turbine blades 140. The deflection limiters 100 also limits the inner turbine duct 128 and the high pressure turbine disk 148 from coming to far apart which would allow hot gas to flow into the cavity located radially inward of the turbine flowpath 27 and heat up the high pressure turbine disk 148 potentially leading to failure.
The exemplary embodiment of the deflection limiters 100 disclosed herein include outward projections 80 extending upwardly or radially outwardly from the middle outer flange 72. The outward projections 80 may be made from two circumferentially adjacent outward complimentary portions 82 on the two circumferentially adjacent singlets 31 in the turbine nozzle assembly 32 as illustrated in
While there have been described herein what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein and, it is therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention. Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims.
