Gas turbine repair

A method for in-field improvement or repair of a stationary gas turbine is carried out by successive removal of the gas turbine discs including its multiplicity of blades (buckets) from the common drive shaft. The overlying ring of tip shoes (shroud) and gas feed nozzles (stationary buckets) are similarly sequentially removed. Each disc is axially and circumferentially index marked with respect to the common drive shaft prior to removal of a portion of the drive shaft and substitution of a dummy drive shaft. The dummy shaft in turn is similarly indexed to the turbine disc index marks. Such marking assures that the turbine discs may then be reassembled to the common drive shaft so that no field rebalancing of the common drive shaft, and its assembled air compressor stages, is required for vibration control. The tip shoes, including the labyrinth knife edge, are then replaced by recoated shoes for control of erosion resistance and progressive abradability of the edges to accommodate centrifugal and thermal expansion of the blade tips. Deposition of the abradable material on the blade tips or erosion of the blades is avoided by reducing the clearance between the tip shoe edges and the blade tips with such progressive hardness of the abradable material on the tip shoes. Such reduction greatly increases efficiency of the engine by enhance of power out put and reduced fuel consumption.

The present invention relates to hot gas turbine engines. More 
particularly, it relates to a method of field repairing, restoring or 
improving clearance between the outer tip ends of the hot gas stage 
turbine blades and the surrounding tip shoes to reduce substantially 
by-passing of hot gases around the ends of the turbine blades so that full 
power is extracted from such gases passing through the blades of several 
turbine discs. The method is particularly directed to a method of 
disassembly and reassembly of only the hot gas section of a gas turbine, 
without removing the entire turbine package from its stationary position 
or from its normally horizontal operating work position. 
BACKGROUND OF THE INVENTION 
Hot gas turbines have been widely used since their inception and first 
introduction approximately fifty years ago. In particular, they are found 
in many stationary applications where a continuous and reliable source of 
power is required, such as to generate of electrical power at remote 
locations, to pump gas or petroleum at high volume for extended periods of 
times, including many months, or even years. They also find many 
applications in transportation such as jet engines, marine propulsion, and 
tractor trailer combinations or big rigs. They vary in size from several 
thousand kilowatt hours per day, as in oil field operations, to generate 
electrical pumping power, or for gas injection and the like tip to many 
megawatts as in electric power generation and distribution. Although hot 
gas turbines are highly efficient overall, and relatively economical to 
operate, they range in cost from $500,000 to a few millions dollars, for 
modestly sized units, such as those that generate a few hundred kilowatts 
per day. Larger stationary hot gas turbine units, generating several 
hundred thousand kilowatt hours per day may cost from several million 
dollars up to several tens of millions of dollars. 
It has long been known that the efficiency of such hot gas turbines is 
highly dependent upon full utilization of power generated by combustion of 
hydrocarbon liquids or gas with high pressured air generated by the air 
compressor section of a gas turbine engine. In its general configuration, 
a hot gas turbine comprises an air compressor section having a 
multiplicity of air compressor stages, that compress air to several 
hundred pounds per square inch, as it successively passes through the 
several compressor stages. The compressor stages are driven as a unit by a 
single common drive shaft, that extends through a combustion zone where 
hot gas is generated by the compressed air and fuel to drive the turbine 
section at the other end of the common drive shaft. The hot gas section is 
made tip of a plurality of hot gas stages each of which includes a turbine 
blade disc directly mounted on the common drive shaft. Power from the 
drive shaft is delivered to either one, or both ends, through suitable 
gearing, to drive any rotatable load, such as an electrical generator, an 
auxiliary pump, or the like. 
While it has been known for a number of years that the efficiency of a hot 
gas turbine can be improved if clearance between the tips of the rotating 
blades on each turbine disc and the surrounding tip shoes can be reduced 
to a minimum so as to prevent hot, high pressured gases from bypassing 
each set of turbine blades by flow around the outer edges or tips of the 
blades. Unfortunately, for greatest efficiency, the rotor blades must be 
operated at high temperatures and high speeds that cause thermal and 
centrifugal forces that radially expand such turbine blades. Additionally, 
gravity, work loads and vibration forces acting on the turbine discs tend 
to cause the blade tips to "carve" or erode away portions of the 
surrounding tip shoes. In particular, the lower portion of a ring of the 
surrounding tip shoes, as viewed from the center line of the rotating 
common drive shaft may permit the blade tips to cut a non-circular path to 
form such an oval track over an arc of from a few degrees to a hundred 
degrees or more of the circumference. 
Because of these changes in path of the blade tips in going from cold, or 
ambient, conditions to normally high operating temperatures of 900.degree. 
to 1850.degree. F. and at high speeds, clearance, or "gaps," between the 
blade tips and the tip shoes can vary widely around the circumferential 
path. To avoid potential damage to the disc blades, or possible 
destruction of the turbine hot stage, and even ballistic destruction of 
the casing or anything external around its circular path, we have found 
that factory or overhaul gaping is frequently too wide to achieve ideal or 
"rated" clearance so that no possible contact will occur between the two 
surfaces. And although abradable materials have been known, both for 
coating the ends of the turbine blades and the tip shoes, such materials 
have generally been used only on the air compressor stages which are not 
subjected to temperature changes of several hundred degrees, as compared 
to the hot stage turbine blades and tip shoes. Furthermore, where such 
abradable material has been added to either the blade tips or the tip 
shoes, the clearance is normally set so that little or no abrasion occurs. 
This appears to be due to a fear that such abraded material from the first 
or second hot stage discs may damage or plug the nozzles of a subsequent 
turbine stage. 
Accordingly, little or no attempt seems to have been made to adjust the 
clearances between the turbine blades and tip shoe after the turbine is 
installed and operating. Thus, the only alternatives to operation at less 
than rated power, has been to keep running at the same low-efficiency, or 
to rent a replacement turbine, or to shut-down and ship the entire engine 
to an overhaul, or factory repair facility. Furthermore, correction of the 
gaping has not always assured that such an overhaul will in fact increase 
the power output of the installed turbine. The costs of shutting down or 
renting a substitute engine, and the cost of shipping and repair of the 
turbine can frequently exceed 100,000 dollars to several hundred thousand 
dollars. Where the gas turbine is located in a remote location, it is of 
course even more costly and more time consuming to have the engine 
packaged, and sent to be overhauled. Hence, there has long been a need for 
a field repair or overhaul procedure for improving the hot gas section of 
a gas turbine. However, there are other reasons which appear to have 
prevented people from attempting such field overhaul. These relate to a 
fear of possible injury of people or equipment, if any of the hot gas 
stages fail mechanically. A particular danger in this regard is that in 
factory assembly or overhaul, it has been considered essential to 
dynamically rebalance the common drive shaft, including both the air 
compressor assembly and the hot gas assembly. Since each of these 
assemblies is made up of a multiplicity of discs and their multiplicity of 
blades, such initial balancing, or subsequent rebalancing present a 
complex vibration problem. That problem relates to the axial spacing of 
and tile two assemblies from each other and their wide range of speeds, as 
well as the flexibility of the common drive shaft. 
The length of tile drive shaft and tile separation of such rotating masses 
on the drive shaft requires careful dynamic balancing of the rotating 
components as a unit, at up to their maximum operating speeds. Such 
balancing is to avoid serious damage of the hot gas turbine blades due to 
vibration at any speeds over which the common drive shaft may operate. 
Since the same drive shaft also is coupled to a rotatable load either at a 
single end or at both ends of the drive shaft, any rotational vibrations 
of the drive shaft present an exceedingly difficult problem to handle. 
For the foregoing reasons it will be understood that although there have 
been great incentives to develop a field reliable method, or procedure, 
for overhauling the hot gas section of a gas turbine engine, such need has 
not been answered prior to our invention. In carrying out our invention, 
we have devised methods, apparatus, and materials, that permit such 
economic rehabilitation of a stationary, or a transport, gas turbine 
engine without danger of modifying the dynamic balance of a turbine drive 
shaft during such repair or overhaul including all of its various 
components at any location, world wide. At the same time such overhaul or 
rehabilitation of the hot gas section is performed without removal of the 
complete gas turbine engine, from its normal outer package or its 
interconnections with its rotatable load. Accordingly, repair and 
installation of modified, and more effective, tip shoes, including 
abradable materials substantially reduce clearance between the tip shoes 
and the turbine blade tips is now possible and thereby substantially 
improves the power output and overall efficiency of the gas turbine 
engine. At the same time, prolonged down time of the turbine installation, 
is significantly reduced to a period of time to not more than a few hours, 
to a day or two. Thus overhaul, for complete disassembly, and assembly, 
including replacement of parts, and re-assembly for return to service, are 
all performed with minimum danger of damage to the turbine or personnel in 
the vicinity of the turbine. 
SUMMARY OF THE INVENTION 
The present invention is particularly characterized by the procedural steps 
whereby the hot gas section of the turbine is disassembled and reassembled 
to carry out an in-field method of rehabilitating or rebuilding the tip 
shoes so that less clearance is required between the tip shoes and the 
blade tips. This is additionally made possible by tailoring the abradable 
material on the face of the tip shoes to assure that it is readily 
removable to a desired gap during installation of the tip shoes. The 
abradable material must not only be highly resistant to heat, but also 
readily abrades by friction contact that may occur if thermal expansion of 
the turbine blade during start-up, or speed excursions, including normal 
vibration of the blade tips, as their path changes in diameter. Further, 
such tip shoe abradable material, if rubbed or worn off, must additionally 
readily exhaust through any subsequent nozzle stages, without damage or 
plugging. 
In accordance with a :first aspect of the present invention, our method of 
on-site disassembly and reassembly of only the hot gas section of the gas 
turbine engine is in part made possible by the conventional construction 
of such turbine engines. As now marketed and used in generating power at a 
substantially constant speed over a period of many thousands of hours, 
such disassembly and reassembly is readily accomplished without removing 
the engine from its normal horizontal operating position because access to 
the hot gas stage is readily available by removal of the exhaust gas 
diffuser housing. Where initial construction of the turbine does not 
readily permit such access to the several hot stages, a one-time 
conversion of the exhaust diffuser housing may be readily effected without 
disturbing any of the rotating elements by a turbine mechanic. Because the 
apex of the conical form of the turbine housing around the hot section is 
near the combustion chamber, the last stage of the hot gas section is 
readily opened. Thus, assembly and disassembly of the hot stages permits 
access around the full circumference of the drive shaft and along the full 
length of the hot gas section. 
With such exhaust gas diffuser open, it is to be noted that the hot gas 
section includes a plurality of hot gas stages, in which the assembly from 
the last stage to the first is in the following order for each such stage 
and includes 1) a turbine disc includes a multiplicity of turbine blades 
that are interlocked around the periphery of the turbine disc, and a hub 
that interlocks with the drive shaft and the adjacent turbine discs, 2) a 
plurality of tip shoe each of which covers an arcuate segments surrounding 
the circumference of the turbine blades and each has a surface that 
closely overlies the full circumferential path of the blade tip ends to 
minimize flow of gas around the outer ends of said blades and 3) means for 
supporting a plurality of stationary gas nozzles, (some in the form of 
turbine blades) for directing a high pressure, combustion gas stream 
against the next adjacent set of turbine blades after passing through the 
previous turbine stage. 
After removing the exhaust gas diffuser housing, at least the lower 
periphery of the whole hot gas section is supported, as a single unit, so 
that each turbine disc, and its surrounding tip shoe segments and nozzles, 
so as to maintain normal, horizontal alignment of the full section with 
the common drive shaft, during subsequent disassembly. 
The next essential step is to index mark the spacial position of each 
turbine disc relative to its axial and circumferential position on said 
drive shaft. Such index marks in part assure that reassembly of the discs 
and drive shaft are not displaced either radially or axially with respect 
to each other. This is the first highly effective step to avoid need to 
rebalance the entire rotating assembly. Then, at least a rearwardly 
extending portion of the drive shaft, on which the said hot turbine stages 
are assembled as a unit, is threadably disconnected and withdrawn without 
rotation between the turbine discs and the balance of the drive shaft. 
A dummy drive shaft that replicates all of the radial and axial dimensions 
of the removed section of the drive shaft, except for the hexagonal drive 
head, is prepared as a substitute. Alternatively, the hex head of the 
original drive shaft may be modified by substituting an internal socket 
for the external hexagonal head. The dummy drive shaft section then 
supports the hot gas stages when substituted by insertion through the said 
rearwardly extending hot gas stages previously mounted on the common drive 
shaft. 
The dummy drive shaft now permits removal of the elements of each hot stage 
be slidable moving of the bore of the hubs of each turbine disc over the 
dummy drive shaft. The similar threaded end secures the dummy drive shaft 
to the forwardly extending portion of the common drive shaft 16 and the 
two sections are preferably torqued together to the same extent as the 
original drive shaft connection between common shaft 16 and section 17. 
The circumferential and axial positions of each turbine disc relative to 
its assembled orientation on the dummy shaft is then registered or indexed 
on the substituted dummy drive shaft. Then successively, each of the hot 
stage elements is removed from the turbine drive shaft over the open end 
of the dummy shaft. The elements are removed in the following order: 
1) the turbine disc, 2) the ring of tip shoes and 3) the corresponding 
nozzle ring. 
Before disassembly of at least the first and second hot stages of the 
plurality of hot stages, the spacial gap between the turbine blade tips 
and the opposed faces of each tip shoe segment is measured and recorded at 
a plurality of spaced apart locations on each tip shoe face. For each tip 
shoe face surrounding the turbine blade tip paths, preferably measurements 
are recorded near their opposite ends and near the center. 
Re-assembly of the multiplicity of hot stages is in the reverse order of 
the above-noted steps. Each turbine disc is oriented around and along the 
dummy shaft by registering its index-marks to the previously recorded 
marks on both the shaft and the turbine discs. However, the opposed faces 
of each replacement tip shoe segment includes an abradable surface coating 
having a Mohs scale hardness of not less than about 60 and not greater 
than about 80. Further, the abradable surface spans the axial width of the 
turbine blades and has a thickness, or depth, extending radially toward 
the turbine blade tip path that is at least 0.060." Desirably, but not 
necessarily, this results in zero gap between the tip shoe face and 
"standard" or cold length of the turbine blades. Thus, if such spacing 
overlaps all clearance space between the shoe tip face coating and the 
opposed blade tips, the excess material is radially abraded on each tip 
shoe face, as by scraping, or grinding, portions of each tip shoe face to 
obtain an average clearance of not more than about 0.040 inch around the 
arc of each tip shoe and around the full circumference of each stage of 
the turbine blade tips. 
After reassembly of the hot gas section, such assembly is again supported 
under its lower periphery, independent of the dummy drive shaft. The dummy 
drive shaft is unthreaded from the turbine drive shaft to permit 
replacement of the original rearwardly extending portion of the common 
drive shaft. Again, each index marker on the original common drive shaft 
section is realigned with the index markers on each turbine disc, both 
radially and axially so that the reassembled discs on the turbine drive 
shaft do not require dynamic rebalancing prior to operation of the 
re-assembled gas turbine. 
Further objects and advantages of the present invention will become 
apparent from the following detailed description of the preferred method 
to carry out the necessary steps of assembly and reassembly of only the 
hot gas section of a gas turbine to restore or enhance the efficiency of 
both fuel consumption and power output and without rebalancing the rotary 
assembly, as described and shown in the accompanying drawings, which form 
an integral part of the application.

PREFERRED EMBODIMENTS OF THE INVENTION 
Referring now to FIG. 1, the method of the present invention is 
particularly directed to restoring or improving the efficiency of hot gas 
turbine engines 10. The primary elements or such a turbine engine are 
shown in a block diagram by FIG. 1, as used for power generation. Such 
engines are most useful at remote or undeveloped geographic locations to 
generate electrical power, or to pump oil or gas, or to drive similar 
rotatable loads. As shown, the primary component of engine 10 include an 
air compressor section 12 comprising a multiplicity of air compressor 
discs (not shown) rotatably mounted on common drive shaft 16 extending 
through a forward portion of turbine housing 18. Housing 18 also encloses 
a combustion chamber 20 wherein fuel oil, or gas, combines with compressed 
air generated by air compressor section 12 to supply hot gas to drive 
turbine discs forming hot gas section 22, similarly rotatably mounted on 
common drive shaft 16. 
Hot gases from combustion chamber 20 are fed to turbine hot gas section 22 
through a group of first stage feed nozzles 28 around first stage turbine 
blades 24 mounted on disc 30, best seen in FIG. 2. Disc 30, along with 
second stage disc 32 and third stage disc 34, jointly drive common drive 
shaft 16. Common drive shaft 16, in addition to driving compressor discs 
14, directly rotates power load units, such as a generator or pump, 
through reduction gearing 36 or 37 at either, or both, ends, of engine 
housing 18. Hot gases exhaust from third stage disc 34 are collected and 
diffused either as exhaust gas, or for heat recovery, by flow into gas 
diffuses housing 38 and out of vent 39. 
As noted above, because it has frequently been found that an engine is 
unable to achieve its ideal design or "rated" power to drive a given load, 
it has been necessary to tolerate less than optimum output, rather than 
incur the expense of renting a substitute gas turbine engine (frequently 
not available, or not economically feasible). Alternatively, total loss of 
use of the engine for many days or weeks may be necessary to complete such 
repairs or overhaul. Accordingly, there has long been a need for a method 
that permits in-field, or on-site, disassembly and repair to recover lost 
power required of such engines, namely by reducing the clearance space 
between the turbine blade tips 25, 62 and 64 and the arcuate faces 49 of 
the multiplicity of tip shoes 50, 52 and 54, surrounding the blade tips. 
In accordance with the present invention, we have developed procedures that 
permits achieving such rated, or upgraded, performance of a gas turbine 
engine, such as engine 10, so that in-field disassembly and reassembly 
does not require removal of the complete engine assembly for shipment to 
an overhaul facilities. Rather, the steps of the present invention permit 
systematic removal of only the elements of hot gas turbine section 22, 
without requiring rebalancing of the drive assembly over its full dynamic 
range of drive shaft speeds. 
Such in-field replacement of portions of hot gas section 22 must be carried 
out without requiring such dynamic rebalancing of the complete drive shaft 
assembly 16, including air compressor section 12. For this to be possible, 
it is essential that each of turbine discs 30, 32 and 34 be registered or 
indexed relative to each other and their precise axial and 
circumferential, or radial, positions on drive shaft 16. As will be 
apparent such registration eliminates the normally required rebalancing of 
the entire drive shaft assembly after disassembly and reassembly of the 
entire turbine engine. Additionally, we have found that removal of exhaust 
gas diffuses 38 readily exposes only the portion of the interior of engine 
housing 18 that substantially surrounds the three hot gas stages drive 
discs 30, 32 and 34. Most importantly, removal of diffuses 38 exposes a 
rearward portion of drive shaft 16 which is formed as a tubular member 17 
that includes a threaded forward end 42 that couples portion 17 to similar 
threads 43 at the rear end of drive shaft 16. The opposite end of shaft 17 
includes a hexagonal head 44 for rotating and axially compressing, or 
torquing, shaft section 17 to shaft 16. When shaft 17 is fully torqued, 
discs 30, 32, and 34 are bound together at their respective hubs 31, 33 
and 35 to form a single rotating mass, or body, integrate with drive shaft 
assembly 16, and the multiplicity of similar discs forming air compressor 
12. It will be noted that the diameter of shaft 17 is stepped so that 
forward end 42 is smaller. Hub 31 of disc 30 has a matching smaller bore 
61 which extends through the forward portion of hub 33 of disc 32. The 
rearward diameter of shaft 17 has a larger diameter that matches bore 63 
passing through the rearward portion of hub 33 and hub 35 of discs 34. 
With exhaust gas diffuser 38 open, the first essential step of in-field 
repair for enhancement of the hot gas section is to index both the axial 
and radial positions of each turbine disc hub 31, 33 and 35 relative to 
drive shaft portion 17 and in turn section 17's radial position relative 
to the rest of common drive shaft 16. As best seen in FIG. 2, flanges 29, 
31, 33 and 35 are formed with interlocking gear teeth or lands 41 around 
shaft 17. Before any disassembly of hot gas section 22, the assembled ends 
are each marked as to their junction points on the opposing and 
interlocking faces of hubs 31, 33 and 35, as well as their positions on 
portion 17 and main drive shaft 16. 
The indexed junction marks are vital because they assure that upon 
reassembly into the same configuration, no radial or axial displacement of 
any of the rotating elements has occurred. Thus, when operating the engine 
over its normal operating ranges of speeds, temperatures and loads, the 
dynamic balance of the total drive shaft assembly will not change. Index 
markings of discs 30, 32, and 34 relative to drive shaft portion 17 may be 
made either before, or after, at least the lower half of the axial length 
of hot gas section 22 is externally supported by a sling either held by 
bracing, or jacks, (not shown) from below the full hot gas assembly or 
from above hot section 22 by suspension from a sling (not shown) supported 
by an overhead crane, or from the apex of a rigid tripod (not shown). 
After suspension of hot gas section 22, threaded end 42 of drive shaft 
portion 17 is unscrewed from drive shaft 16 by rotation of hex head 44, 
preferably using a torque wrench. Additionally advantage may be taken of 
the fact that shaft 17 is generally annular so that its interior is 
tubular. This permits heating of the threaded connection by an electric 
heater element having an external diameter that allows insertion of the 
heater to a depth sufficient to thermally expand threads 42 to disconnect 
shaft 17 from shaft 16. Preferably, the torque between threads 42 and 
mating threads 43 of drive shaft 16 is recorded and used in reassembly of 
shafts 16 and 17. 
If after removal of shaft section 17, no similar hot gas turbine section 
has been field-disassembled, a dummy shaft 46 must be fabricated. Such 
dummy shaft comprises a replica of shaft 17 so that shaft 46 including 
threads 47 has precisely the same step diameters over the same axial 
lengths as those forming shaft 17. These same lengths and diameters along 
the length of shaft 46 will thus be identical to shaft 17, but without a 
counterpart of hex head 44 (which would prevent axial movement of hub, or 
flanges, 3i, 33 or 35 along shaft 46.) 
Dummy shaft 46, so constructed, is then inserted through disc hubs 31, 33 
and 35 and threaded to shaft 16. Preferably, threads 47 of dummy shaft 46 
are torqued to the same value as threads 42 when coupled to threads 43 of 
drive shaft 16. In the same manner as on shaft section 17, dummy shaft 46 
is indexed to each of the three turbine discs. This procedure is also 
essential to maintain the integrity of all the rotating elements of the 
hot gas section namely the original drive shaft, the air-compressor discs 
and hot gas turbine discs. During such substitution of shaft 46 for shaft 
17, it is essential that no relative rotation occurs between any of hot 
stage discs 30, 32 and 34 and common drive shaft 16. 
Disassembly of the elements of each hot gas stage then begins by removing 
third stage turbine disc 34 by sliding hub 35 along and over the open end 
of dummy shaft 46. However, prior to such removal the gap between each 
blade tip 64 around the circumference of overlying tip shoe 54 is 
measured. Because third stage blades 74 normally operate at lower 
temperatures and pressures, due to conversion of hot gas energy to power 
in passing through the two preceding hot stages discs 30 and 32, tip shoe 
54 may be a single circular hoop or ring, normally having a single narrow 
circumferential gap to permit insertion into step ring 96 around the inner 
surface of casing 18. Further, because of the less sever operating 
conditions, damage is generally less likely to be found on either blade 
tips 64 or face 55 of tip shoe 54. Accordingly, "gaping" (measuring the 
gap) between tip shoe 54 and the blade tip 64 is less likely to require 
change. However, to attain maximum improvement from third stage disc 34, 
the gaps should be recorded around the circumference of the blade tips to 
determine whether the circular path of the third stage blade tips has 
become oval shaped, so that if the amount of gap has increased, its 
clearance over such arc span can be accomodated. 
After removing tip shoe ring 54, the assembly of third stage nozzles 84 may 
be removed as a unit. It will be noted that nozzles 84 are stationary 
relative to disc blades 72 and 74, but each nozzle has a configuration 
similar to those disc blades lying on either side of nozzles 84. The 
multiplicity of nozzles 84 are supported at their outer ends by hangers 
94, also supported by turbine casing, or housing, 18. Hangers 94 are also 
held in place circumferentially by retainer ring 100. It will also be seen 
that the opposite ends 101 of nozzles 84 are sealed against flange 33 and 
35 of discs 32 and 34 respectively by rotary seals 102. Seals 102 prevent 
gas from by-passing the lower ends of nozzles 84. The multiplicity of 
nozzles 84 are desirably removed as a group by an elastic belt (not shown) 
extending around their upper edges, as they are pulled rearwardly away 
from hanger 94. 
Second stage turbine disc 32 and tip shoes 52 are exposed by removal of 
third stage nozzle ring 84. The gaps between tips 62 of turbine blade 72 
and each of the plurality of tip shoe of segments 52 are measured and 
recorded. Preferably, such measurements are made adjacent both ends of the 
arcuate segment forming the tips shoe ring and at the center of each tip 
shoe. In general, eleven tip shoe segments 52, as in FIG. 2, cover the 
circumferential path of blade tips 62. After all gap measurements are 
completed tip shoes 52 are disengaged from step frame 19 of housing 18. 
Tip shoe segments 52 are locked circumferentially by locating pins that 
match a hole in the upper surface of each tip shoe (not shown). Hub 33 of 
turbine disc 32 is then free to slide over the open end of dummy shaft 46. 
Second stage nozzles 82 including hub end 104 and bearing 105 are released 
from frame 19 in the same manner as third stage nozzles 84. 
With the ring of second stage nozzles 82 removed, first stage turbine disc 
30 and tip shoes 50 are exposed. Again, the gaps between each of the 
several arcuate segments forming tip shoes 50 and tips 25 of blades 24 are 
measured. As with the second stage, the gaps at both ends and at the 
central portion of each tip shoe 50 are preferably recorded. Following 
such recording, tip shoes 50 are removed from around blade tips 24 by 
release from circular step frame 19 of housing 18. Turbine disc 30 is then 
free for removal by sliding hub 31 over the open end of dummy drive shaft 
46. If desired, first stage nozzles 28 may then be removed for inspection, 
and if needed, repaired or replaced. 
Based on knowledge that where all engine is not delivering its "rated" 
power, the gap spacing between each tip shoe and blade tips, is generally 
excessive. Such excessive clearance may be due either to there being no 
abradable material between the labyrinth edges of the tip shoes (e.g. edge 
55 of tip shoe 54) or the tip shoes have inadequate abradable coating, as 
by erosion of such abradable material from the labyrinth edges. Excessive 
gaping, or spacing, may also be due either to blade tip damage or to 
partial destruction of the tip shoes after thermal or centrifugal 
expansion of the turbine blade tips that over expands the radial length of 
the turbine blades. Other causes for excess gaps developing may be due to 
gravity sag, bearing wear or excessive vibration developed by the rotating 
drive shaft assembly. Further, such excessive clearance may also be due to 
deliberate cautions as to initial spacing used to avoid any potential 
interference between the blade tips and the tip shoe faces. Other 
potential reasons for excessive gaping include improper installation of 
abradable surfaces material either on the blade tips or on the tip shoe 
faces, or the hardness and/or density of the abradable material, may cause 
failure in normal service. In practice, we have found that due to the 
unpredictable amount and nature of the resulting excess radial clearance 
between blade tips and the tip shoe, the abradable coating, requires close 
control of the width and thickness of the coating over the tip shoe 
surface as well as the support area for the coating on the metal base of 
the tip shoe. Additionally, the density and porosity of the abradable 
coating material as applied by properly relating the tip shoes is 
particularly effective to minimize hot gas loss through such gaps. Minimum 
gap around at least the first and second stage turbine blades is 
particularly effective to increase power output while at the same time 
reducing fuel consumption. Both are highly desirable. To this end, and 
prior to reassembly of the hot gas section, either new or repairable tip 
shoes are preferably prepared and coated in the following manner. 
As shown in FIG. 4A, face 125 of each tip shoe 124 is milled or ground, to 
remove any previously present abradable and corrosion inhibiting material, 
so as to form a smooth cylindrical surface 125. Surface 125 has a reduced 
thickness which terminate at the edges in slots 120 and 122 adjacent rims 
121 and 128, respectively. Together the slots and rims serve to anchor the 
abradable and erosion resistant material filling the space between tip 
shoe face 125 and the turbine blade such as tips 25 or 62 and if desired, 
tips 66. As noted before, such spray material is desirably a mixture of 
nickel-aluminum powder, or nickel, chromium alumina/bentonite component 
powder. Most preferably, we have found that Metco 450-NS Nickel-alumina 
composite powder that is vaporized in a thermo spray gun using 
oxy-actylene is most satisfactory to provide both erosion resistance at 
the base of surface 125 and to form more porous abradable material of 
ridges or knife edges 122, as shown in FIG. 4B. Such choice of the two 
services is provided by reducing the temperature of the sprayed powder to 
form abradable edges 127. The total depth from edges 127 to the base 
surface of face 125 is preferably about 0.060 inches greater than the OEM 
set space between, for example, as in FIG. 2, blade edge 25 and the 
labyrinth edges 49 of tip shoe 50. 
Such added depth is generally adequate to close the space between the edge 
127 of tip shoe 124 to touch the turbine blade tips, at ambient and static 
conditions, unless the blade tips have been eroded, or ground, away. With 
such depth, as each tip shoes 50 is inserted into its locked position in 
hanger 19, sufficient abradable material is available so that it may be 
readily ground, or abraded away, as by emery cloth, or a portable grinder, 
to create a uniform gap clearance of about 0.040 inch between edges 122 of 
face 125 at all portions around the circumference of tips 25 or 62 of 
blades 24 or 82, as shown in FIG. 4C. Although a portion of this circular 
closure around discs 30 or 32 may exceed that optin-ram value over a few 
degrees of arc, (e.g., due to gravity deformation of drive shaft 16) the 
overall gap area for gas loss is substantially reduced. Thus, the total 
volume of hot gas flow driving blades 24 and 72 substantially decreases 
fuel consumption and increases power output. The abradable material 
preferably has a Rockwell hardness of from 40 to 80 so that said material 
will stand up under hot gas temperatures of up to about 1660.degree. F., 
but the abradable material has a porosity such that only small shavings, 
or particles, are cut or abraded away from ridges 122 of tip shoes 50 by 
thermal or centrifugal growth, or expansion, of turbine blade tips 25, but 
without deposition, or blockage, of any of the down stream stages of the 
hot gas section, such as the second or third stage nozzles 82 and 84 or 
their turbine blades 62 and 64. 
Using the specially rebuilt tip shoes in at least each of the first two hot 
gas stages, reassembly of the hot gas section is in the reverse order to 
those described above, for disassembly of the hot gas section. However at 
this stage of the overhaul process, the installation of each of turbine 
discs 30, 32 and 34 is in accordance with its respective registration, or 
index marking, with respect to the adjacent turbine discs and its axially 
and circumferential position on dummy drive shaft 47. These must be 
carefully observed and followed to avoid disturbance of the dynamic 
balance of the complete drive shaft assembly to avoid destructive 
vibration problems in reassembly of the turbine engine, without removal of 
the engine from its normal location. 
After assembly of the hot gas section on dummy drive shaft 47, the dummy 
drive shaft is unthreaded from main drive shaft 16 and original drive 
shaft 17 is reinstalled with the same caution again being exercised to 
register correctly each index mark in the final assembly. Thus, the radial 
and axial positions of each rotatable element, namely, discs 30, 32 and 
34, are interlocked by their hubs 31, 33 and 34 with shaft hubs 29 on 
drive shaft 16, including hex head 44 of hot gas stage shaft 17. Heating 
of the connecting threads 42 and 43 and by tightening shaft 17 to its 
original torques value assures that the full hot gas section 22 is 
completely reassembled without displacement of the engine from its initial 
work position. 
Reassembly of exhaust gas diffuser to enclose section 22 completes 
reassembly of the engine for return to service. 
Further changes and modifications of the present invention will occur to 
those skilled in the art of hot gas turbines from the foregoing 
description and drawing. All such changes, modifications or alterations 
covered by the attached claims are intended to be covered thereby.