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S, T55-L-714 ENGINEDEVELOPMENT AND QUALIFICATION,

TIP CLEARANCE ANALYSIS,LYC-T55 -88-04

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T55-L-714 ENGINEDEVELOPMENT AND QUALIFICATION

CDRL A006TIP CLEARANCE ANALYSIS

LYC T55 88-04

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31 MAY 1988

Prepared by: -..-- ",-

S. Sawdye

Developmen t Engineer

Approved by: - - - ---------------

Dr. H. Jen /Turbine Coofling Manager 0

TABLE OF CONTLNTS

PAGE 0

LIST OF FIGURES .............................. ii

LIST OF TABLES ............................... iii

1.0 OBJECTIVE .............................. 12.0 ENGINE CONFIGURATION................................1I3.0 ANIALYSIS METHOD................................... 14.0 RESULTS........................................ 2

4.1 CYLINDER TEMPERATURES .................... 2

4.2 COMPONENT GROWTH .................... 24.3 TIP CLEARANCE .................... 3

5.0 CONCLUSIONS AND RECOMMENDATIONS ................... 3

6.0 REFERENCES . .............................. 4

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LIST OF FIGURES

FIGURE 1 Gas Producer Turbine Cooling Air Network

2 Gas Producer Turbine Rotor and Cooled Cylinder Assemuly3 First Stage GP Turbine Cylinders (uncooled and cooled)4 Second Stage GP Turbine Cylinders (uncooled and cooled)5 Engine Operating Cycle6 Principal Engine Performance Parameters on Accel and Decel

7 First and Second Stage GP Turbine Cylinder MetalTemperatures at Take-off Steady State

8 First GP Turbine Component Radial Growths9 Second GP Turbine Component Radial Growths10 First GP Turbine Tip Clearance11 Second GP Turbine Tip Clearance

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LIST OF TABLES '

ITABLE 1 First Stage GP Turbine Tip Clearance Summary2 Second Stage GP Turbine Tip Clearance Summary

I1

.0 OBJECTIVE

The performance of the T55-L-714 engine has shown to be strongly influencedby gas turbine running tip clearance. In order to ensure both optimum steadystate performance and to prevent rubs of the gas producer turbine upon rapiddeceleration and acceleration, a steady state and transient tip clearanceanalysis has been conducted. ;,, . , , -

joi2.0 ENGINE CONFIGURATION

The gas producer turbine section of the T55-L-714 engine is shown in Figure1. The cooling flow network includes the reduced disc bore cooling designdescribed in References 1 and 2. The rotor and cooled cylinders are shown inFigure 2.

The cooled 1st stage GP cylinder is shown in Figure 3. Cooling air (0.5%) ofcore flow enters through a number of slotted passages on the cylinder flange.Initially the air flows radially inboard and then forward to the cylinderleading edge where it enters the labyrinth flow passages. The air in thelabyrinth passages flows mainly in the circumferential direction in thegrooves formed between the barriers, as it proceeds from passage to passageto the cylinder trailing edge. At the cylinder trailing edge, the airdischarges and enters the gas stream as film air on the 2nd nozzle outershroud.

The second stage GP cooled cylinder is shown in Figure 4. Cylinder coolingair (2.4%) enters through a series of holes on the cylinder outside wall nearthe flange location. The air flows initially downstream and then upstreamfollowing a flow path which wraps around the cooling insert. The air leavesthe cylinder near its leading edge and flows into the second stage nozzlevane leading edge cooling flow passage for vane cooling. Since the 2.4% ofcooling air for the second stage cylinder is the same air used for the nozzlevane cooling, this airflow is increased by only 0.7% over the current 1.7%flow for cooled nozzle in the -712, yielding a total of 2.4% flow in thecooled cylinder and nozzle configuration.

3.0 ANALYSIS METHOD

The turbine disc, seal plate, rotor blade, and cooled cylinder have beenanalyzed for their growths and effects on the turbine tip clearance underworst engine steady state and transient conditions. The assumed operatingcycle in shown in Figure 5. Engine parameters during acceleration anddeceleration are shown in Figure 6.

A mathematical (nodal) model representing engine geometry was constructedfrom engine drawings. Material nodes were established to represent thecylinders, turbine blades, discs and seal plate. Fluid nodes were added todescribe the boundary conditions surrounding the turbine wheel and cylinderfor both stages. Transient conditions of flow rate, gas temperature andcoolant temperature were imposed on the fluid nodes.

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A computerized method (Ref. 3) was used to perform the mathematicalsimulation. The analytical method is described in (Ref. 5) where tipclearance measurements were made with an optical device and favorablycompared with calculated predictions. Using this method, temperaturedistributions were computed for each node based on the transient and steadystate conditions. Thermal and centrifugal growths were established in theradial direction and resultant tip clearances determined.

4.0 RESULTS

4.1 CYLINDER TEMPERATURES

Cylinder temperatures were computed for all operating conditions and areshown at sea level, take off (maximum power), 130'F day for the worstengine operating condition. Average circumferential thermal conditions areused in the analysis.

The first stage GP cooled cylinder metal temperatures are shown in Figure7(a). The -714 cooled cylinder has an average metal temperature of 1480*F,whereas, the -712 average metal temperature for the uncooled cylinder was1750 0 F.

The second stage GP cooled cylinder metal temperatures are shown in Figure7(b). The -714 cooled cylinder has an average metal temperature of 1580°Fwhile the -712 uncooled cylinder had an average metal temperature of 1670°F.

4.2 COMPONENT GROWTHS

The radial growth in the turbine stage can be described in terms of thefollowing components: disc, serration (eg. disc, rim and blade shank), bladeairfoil and cylinder. Plots of component growths for the first and second GPturbine stages are shown in Figures 8 & 9. Transient conditions with a steadystate stabilization were evaluated: a) steady state ground idle b) groundidle to take off c) steady state maximum power and d) take off to groundidle. Discussion of the growth characteristics for each of the componentsfollows.

The centrifugal growth of the rotating system is instantaneous with the speedchange. The resulting elastic growth of the rotor system is maximum athighest speed. The thermal response of the components is dependant on therate of gas or air temperature change surrounding the component as well asthe thermal lag of the component. Thermal lag is increased with componentmass and to a lesser extent is reduced by high heat transfer rates, whichoccur when exposed to large heat transfer coefficients.

The disc accounts for a major portion of the rotating system growth. Duringtransients the large mass of disc compared with the other components causesit to lag in thermal growth. This causes the exponential decay shape of itstransient growth plot, which typically stabilizes after several minutes. -

2

The serration (disc rim and blade shank) growth is small since the radiallength of this area is small. Although this area is moderate in mass, itstill has a lag in thermal growth due to small heat transfer coefficients andrelatively good thermal contact with the disc. Thus, its growth rate closelyfollows that of the disc. The serration growth is seen as a curve slightlyabove the disc curve in the plots.

The blade experiences the other major portion of the rotating system growth.The blade centrifugal and thermal response is rapid; the blade thermalchange occurs in less than 10 seconds.

Cylinder growth is due only to thermal growth. The purpose of cooling the '

Cylinder is to reduce its growth during all of the engine operation; thereby"%reducing tip clearance. The transient response of a cooled cylinder isslower than that of the uncooled one, because the air cooling heat transfercoefficient is higher in the cooled cylinder. In the T55-L-714, thecylinder's response rate is intermediate between that of the blade and disc.

In general, the observations from Figures 8 and 9 are similar and follow theabove discussion. Minimum clearance occurs during transient conditions.

4.3 TIP CLEARANCE S

The differential growth of the cylinder minus the blade tip is shown inFigures 10 and 11 for the first and second GP turbine stages respectively.This represents "relative" tip clearance assuming zero assembly clearance.Allowances for eccentricity, rotor blade run-out and cylinder distortionunder hot running conditions are routinely made to avoid blade tip rubs.Through engine test experience with the cooled cylinder, it appears thatapproximately 0.020 inch must be allowed to avoid local distortion in thehardware. To obtain this value, an assembly clearance of 0.021 inch and0.022 inch is required for the first and second GP turbines respectively.

Two mission profiles were considered for the tip clearance analyses: a sealevel standard 59*F and 130 0 F day. As expected, tighter tip clearances werefound to exist for the 59*F day transient profile due to cooler compressorexit temperature and subsequent lower cylinder metal temperatures. Relative,assembly, and operating tip clearances are summarized in Tables I and II forthe first and second GP turbine stages. The minimum clearance for theT55-L-714 first GP cooled cylinder occurs during deceleration from take off,steady state to the ground idle power condition. Minimum tip clearance forthe second GP cylinder occurs during acceleration from steady state, groundidle to take off, maximum power.

An engine performance benefit is expected based on steady state operating tipclearance reduction due to redesign from the -712 uncooled to the -714 cooledcylinder configuration. The first stage GP turbine tip clearance reductionfrom the -712 to the -714 was from 0.079" to 0.031" respectively. This isequivalent to 3.5% of blade height. The second stage GP turbine tipclearance reduction was from 0.057" to 0.027", which is equivalent to 1.7% ofblade height.

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5.0 CONCLUSIONS

Using the cooled cylinder design in the -714 engine will improve its

performance due to the following analytically determined advantages:

o Lower cylinder metal temperatures.

o Improved roundness of the cylinder.

o Smaller tip clearances.

Final verification that the cooled cylinder design will enhance engine

performance and ensure proper tip clearance control will be accomplished

during qualification testing. A,

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6.0 REFERENCES

1. L. Hayes, "T55-L-712 First GP Turbine Disk Temperature Analysis with

Reduced Bore Cooling Air", Report No. LYC 81-29, March 1981.

2. "T55-L-712 Performance Improvement via Leakage Reduction", Report No.

LYC 82-4, January 1982.

3. E. Hartel, C. Duhaime, W. Corcoran, "Transient Tip Clearance Modeling

Program" Lycoming Computer Program F119A, revised March 12, 1980.

4. S. White, "Turbine Tip Clearance Measurement", Report No.USARTL-TR-77-47, Dec. 1977. Final report for AVRADCOM Ft. Eustis, VA

prepared by Avco Lycoming Division.

5. A. Chandoke, "T55-L-712: Tip Clearance Analysis of the First and SecondGas Producer Turbine Assembly with Uncooled and Cooled Cylinders without

Abradable Coating.", Report No. LYC 83-39, July 1983.

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