Concentricity ring

There is provided a method of assembling a gas turbine engine using a concentricity ring to compensate for assembly tolerances. The blade tips of the gas turbine during turbine rotation define a tip surface of rotation concentric the shaft axis. Due to cumulative tolerances in part manufacture and assembly, the shaft axis is invariably radially eccentric the engine axis at the axial position of the turbine an assembly eccentricity within the range between zero and a predetermined allowable assembly tolerance. The method relates to the step of positioning a concentricity ring with: an outer cylindrical surface engaging the internal cylindrical surface of the engine housing; and an inner cylindrical surface engaging the external cylindrical surface of the turbine shroud, the outer and inner cylindrical surfaces of the concentricity ring being eccentric a ring eccentricity within the range between zero and said predetermined allowable assembly tolerance to rectify the assembly eccentricity of the moving and static components in a simple expeditious manner.

TECHNICAL FIELD
 The invention is directed to a method of assembling a gas turbine engine
 using a concentricity ring between mating components to compensate for
 cumulative assembly tolerances, and in particular to achieve an extremely
 close tolerance which is essential in assembling high pressure gas
 turbines within a mating turbine shroud to minimize the blade tip gap and
 optimize engine performance.
 BACKGROUND OF THE ART
 In the assembling of a gas turbine engine several interconnecting
 components are assembled together, all of which are machined with high
 precision although within a specified dimensional tolerance. The
 cumulative effect of these tolerances in assemblies of multiple parts must
 be accommodated especially when extremely high precision in assembly is
 required.
 One example of such precision is in the assembly of high-pressure turbines
 and their associated turbine shrouds immediately downstream of a combuster
 in the gas turbine engine. It has been estimated that the tip clearance,
 between the tips of turbine blades and the turbine shroud surrounding the
 rotating blades, is so critical to engine performance that each 0.001-inch
 of excess tip gap results in approximately a 0.25% decrease in engine
 performance. As a result, extreme care is taken in ensuring that
 high-pressure turbine blades are accurately assembled together with their
 associated turbine shrouds.
 In the prior art the method of assembling high-pressure turbines and their
 shrouds involves assembling the engine including the combuster and
 high-pressure turbine shaft within acceptable assembly tolerances. The
 high-pressure turbines are then mounted to the high-pressure turbine shaft
 and invariably the rotational axis of the turbine is somewhat eccentric of
 the longitudinal axis of the combuster due to accumulation of machining
 tolerances.
 In order to accurately fit the turbine shroud about the blades of the
 turbine with minimal tip gap, it is common practice to custom grind the
 internal surface of each shroud to precisely match the eccentricity and
 outside diameter of the turbine blades in operation.
 As will be appreciated, the withdrawal of the turbine shroud from assembly
 operations and precision grinding involve significant delay in the
 assembly operation as well as high costs as a result of the skilled labour
 involved in the process. In effect, while the custom grinding operation is
 being carried out on the turbine shrouds, the assembly operation is halted
 imposing significant manufacturing difficulties due to space and
 scheduling commitments.
 In general, all machinery assembly operations can benefit from the use of
 standardized components rather than custom fitting each component as
 required. Standardization of components and simplification of the assembly
 process, inevitably will reduce costs and increase the speed of
 production.
 It is an object of the present invention, therefore, to eliminate the
 custom grinding of turbine shrouds and permit the standardization of
 turbine shroud manufacture regardless of the eccentricity produced in
 assembly of turbines used in gas turbine engines.
 It is a further object of the invention to provide a simple means by which
 eccentricity of turbines can be accommodated without custom grinding or
 removal of components from the flow of assembly and production.
 It is a further object of the invention to enable eccentricity of the
 assembled turbines to be accommodated without significant alteration to
 the prior art structure of gas turbine engine components. Ideally, only
 minimal modification is desirable.
 DISCLOSURE OF THE INVENTION
 The invention provides a method of assembling a gas turbine engine using an
 eccentic concentricity ring to compensate for assembly tolerances.
 A typical gas turbine engine has a longitudinal engine axis with at least
 one turbine disposed on a shaft assembly at a predetermined axial position
 for rotation about a longitudinal shaft axis. The turbine includes a hub
 with a circumferentially spaced apart array of turbine blades each having
 a radially outward blade tip. The blade tips during turbine rotation
 define a tip surface of rotation concentric the shaft axis.
 Due to cumulative tolerances in part manufacture and assembly, the shaft
 axis is invariably radially eccentric the engine axis at the axial
 position of the turbine an assembly eccentricity within the range between
 zero and a predetermined allowable assembly tolerance. The engine is
 assembled progressively to the stage where the static outer engine housing
 assembly has an engine housing flange with a cylindrical internal surface
 concentric the engine axis, and a static turbine shroud having a shroud
 flange with an external cylindrical surface is to be removably mounted to
 the engine housing flange with an internal turbine shroud surface of
 rotation matching the blade tip surface of rotation.
 The invention relates to the step of positioning a concentricity ring with:
 an outer cylindrical surface engaging the internal cylindrical surface of
 the engine housing flange; and an inner cylindrical surface engaging the
 external cylindrical surface of the turbine shroud flange, the outer and
 inner cylindrical surfaces of the concentricity ring being eccentric a
 ring eccentricity within the range between zero and said predetermined
 allowable assembly tolerance to rectify the assembly eccentricity in a
 simple expeditious manner.
 Further details of the invention and its advantages will be apparent from
 the detailed description and drawings included below.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS
 With reference to FIG. 1, the invention is directed to a method of
 assembling a gas turbine engine using a concentricity ring 1 between
 mating cylindrical components to compensate for cumulative assembly
 tolerances and machining tolerances of the assembled engine components. In
 particular the invention is directed to achieving the extremely close
 tolerance essential for assembling high pressure gas turbines within a
 mating turbine shroud to minimize the blade tip gap and optimize engine
 performance. The general construction of a gas turbine engine will be
 considered part of the common general knowledge to those skilled in the
 art and therefore will be briefly outlined in this description.
 As shown in FIG. 1, a gas turbine engine includes a longitudinal engine
 axis. In the embodiment illustrated, the engine includes two high-pressure
 turbines 3 disposed in a shaft assembly 4 at a pre-determined axial
 position for rotation about a longitudinal shaft axis 5. The shaft axis 5
 is shown in FIGS. 1 and 3 as eccentric from the engine axis 2 by a
 dimension "e". It will be understood that for clarity the drawing is not
 to scale and that in reality the eccentricity dimension "e" is in the
 order of less than 5/1000.sup.th of an inch which would be imperceptible
 in the scale of the drawing attached.
 It will also be understood that although the description relates to
 assembly of a high-pressure turbine, the invention is equally applicable
 to any machine or turbine assembly where eccentricity is to be compensated
 for. A high pressure turbine is merely chosen for this example since it is
 an area of the gas turbine engine in which extremely close tolerances are
 absolutely essential and where excessive blade tip gap can result in
 extremely high engine efficiency losses.
 Referring jointly to FIGS. 1 and 2, a typical gas turbine engine includes
 turbines 3 which have a central hub 6 and a circumferentially spaced apart
 array of turbine blades 7 each having a radially outward blade tip 8. It
 will be apparent that during turbine rotation the blade tips 8 define a
 tip surface of rotation which is concentric to the shaft axis 5. Each
 blade 7 will be machined within the appropriate machining tolerance
 however; each individual blade tip surface will be slightly different. The
 rotating blade tips 8 therefore define a tip surface of rotation, which
 represents the radially outward most extent of any blade 7 within the
 array. Of course, ideally all blade tips 8 are identical and define a
 generally frustoconical tip surface of rotation concentric to the shaft
 axis 5.
 The shaft axis 5 is radially eccentric the engine axis 2 at the axial
 position of the turbines 3 by an assembly eccentricity "e" which is within
 the range between 0 and a pre-determined allowable assembly tolerance. For
 instance, in a relatively small diameter engine, the turbine diameter will
 be in the range of 20 to 30 inches and the eccentricity allowable on
 assembly will be in the order 5 to 10/1000th of an inch.
 In assembling the gas turbine engine illustrated in FIGS. 2 and 3, the
 progressions of assembly is generally from left to right as shown. The
 stage of assembly relevant to the present embodiment of the invention is
 reached when the shaft assembly 4 is completed absent the high-pressure
 turbines 3 and the cover plate 9, the combustor 10 is completely assembled
 with stator assembly 11 up to bolts 12, and the combustor outer case 13 as
 a static outer engine housing assembly is completed up to the engine
 housing flange 14. With the left to right assembly of the engine completed
 to this stage, the engine is ready for the next step of installing the
 turbines 3 on the shaft assembly 4 and simultaneous assembly of the static
 turbine shroud 15 with a shroud flange 16 removably mounted to the engine
 housing flange 14.
 As mentioned above, prior art assembly methods do not include the
 concentricity ring 1 nor the particular arrangement of the engine housing
 flange 14 and shroud flange 16. Prior art assembly of the static turbine
 shroud 15 to the high-pressure turbines 3 involves precise custom grinding
 of the static turbine shroud 15 to suit the blade tip surfaces of rotation
 for each engine configuration on assembly. Referring to FIG. 2, prior art
 systems of assembly include the mounting of turbine 3 on the shaft
 assembly 4 and measuring the eccentricity "e" for each engine assembly.
 The turbine shroud 15 as shown in FIG. 2 immediately radially adjacent the
 blade tips 8 includes a separate annular ring 17 which in prior art
 methods is initially prepared with an internal diameter less than the
 external diameter of the tip surface of rotation in order to accommodate
 the custom grinding operation. The internal surfaces of the annular ring
 17 in the prior art is custom ground to match the eccentricity "e" and to
 maintain close blade tip gap tolerances. It will be appreciated as
 described above that this custom grinding operation severely delays the
 engine assembly operation and involves intensive highly skilled labour.
 In contrast the present invention permits the annular ring 17 to be
 precisely machined on their interior surfaces to the desired configuration
 completely prior to the assembly operation. The eccentricity "e" between
 the engine axis 2 and assembled shaft axis 5 is accommodated by the use a
 concentricity ring 1 as will be explained below. Through use of the
 concentricity ring 1 custom grinding of the annular rings 17 and the
 accompanying delays in engine assembly and increased labour can be
 completely avoided. In accordance with the invention the static outer
 engine housing assembling 13 includes an engine housing flange 14 that has
 a cylindrical internal surface 18 concentric to the engine axis 2. The
 static turbine shroud 15 includes a shroud flange 16 with an external
 cylindrical surface 19. The shroud flange 16 is removably mounted to the
 engine housing flange 14 with bolts (not shown) through aligned bolt holes
 20. The annular rings 17 comprising part of the static turbine shroud 15
 have internal turbine shroud surfaces of rotation 21 matching the blade
 tip surface of rotation. A significant advantage of the invention is that
 the blade tip surface of rotation can be accurately determined by precise
 machining of the blade tips 8 and the internal turbine shroud surface of
 rotation 21 can be pre-manufactured to the precise configuration required
 to maintain blade tip gap at a desired level.
 Radial adjustment for maintaining concentricity of the static turbine
 shroud 15 and turbines 3 is provided through the use of a concentricity
 ring 1 as follows. The concentricity ring 1 is precisely machined with an
 outer cylindrical surface mating the internal cylindrical surface 18 of
 the engine housing flange 20. The concentricity ring 1 has an eccentric
 inner cylindrical surface which mates the external cylindrical surface 19
 of the turbine shroud flange 16.
 The outer and inner cylindrical surfaces of the concentricity ring 1 are
 eccentric a ring eccentricity dimension within the range between 0 and the
 pre-determined allowable assembly tolerance. Therefore by positioning the
 concentricity ring 1 between the flange cylindrical surfaces 18 and 19 the
 radial position of the turbine shroud 15 can be adjusted to suit the
 corresponding radial position of the turbine 3 rotating above the shaft
 axis 5. The bolt holes 20 are oversized to accommodate for the outset
 positioning of the turbine shroud 15 as a result of the insertion of the
 concentricity ring 1. Alternatively the bolt holes 20 can be initially
 machined undersized and then are reamed on final assembly.
 It will be understood that different engines will be assembled and result
 in eccentricity dimensions "e". The eccentricity present in the
 concentricity rings 1 applied to different engines may vary anywhere
 between 0 eccentricity and to a pre-determined allowable assembly
 tolerance, for example in the range of 5 to 10/1000.sup.th of an inch. To
 accommodate the expected variance in the eccentricity dimension "e", the
 invention also includes a kit for engine assembly, with a number of
 concentricity rings, 1 wherein each ring of the kit has a different ring
 eccentricity. For example the engine assembly kit can include
 concentricity rings 1 of progressive equally stepped ring eccentricities
 beginning at 0 eccentricity and at each 1/1000.sup.th of an inch to a
 maximum allowable assembly tolerance, of 10/1000.sup.th of an inch for
 example.
 In addition since the eccentricity of the concentricity ring 1 and the
 eccentricity of the high pressure turbines 3 must correspond, the rings 1
 preferably include markings on a ring at various positions corresponding
 to the radial ring thickness at maximum ring thickness, minimum ring
 thickness and medium ring thickness for example. In this way the installer
 can correctly aligned the eccentricity of the concentricity ring 1 with
 the eccentricity of the engine shaft assembly 4 and turbines 3 mounted
 thereon.
 Although the above description and accompanying in drawings relate to a
 specific preferred embodiment as presently contemplated by the inventors,
 it will be understood that the invention in its broad aspect includes
 mechanical and functional equivalents of the elements described and
 illustrated.