NASA

Technical Memorandum 107448

/iJ- cJ 7

Army Research Laboratory

Technical Report ARL-TR-1397

JTAGG II Brush Seal Test Results

Gul K. Arora

AlliedSignal Engines

Phoenix, Arizona

and

Margaret P. Proctor

Lewis Research Center

Cleveland, Ohio

i

Prepared for the

33rd Joint Propulsion Conference and Exhibit

cosponsored by AIAA, ASME, SAE, and ASEE

Seattle, Washington, July 6-9, 1997

National Aeronauticsand

Space Administration

U.S. ARMY

RESEARCH LABORATORY

https://ntrs.nasa.gov/search.jsp?R=19970025652 2018-05-11T16:33:03+00:00Z

JTAGG II BRUSH SEAL TEST RESULTS

Gul K. Arora

AlliedSignal Engines

Phoenix, Arizona 85072-2181

and

Margaret P. Proctor

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135

Abstract

The Tri-services JTAGG II engine uses two identical

brush seals, in tandem, located aft of the high pressure

compressor. The engine operating conditions, at

intermediate rated power (IRP), for this seal are estimated

to be 50 000 rpm (899 ft/sec) speed, 175 psid air to air

pressure differential and 1200 °F air temperature. The

testing was comprised of static air leakage, performance,

seal offset, rotor run out tests and a 50 hr endurance test in

the NASA Lewis seal rig. Based on the test results, it is

concluded that the brush seal design should be able to meet

the air leakage flow factor goal of less than 0.004 for the

engine IRP operating conditions. As a comparison, a

labyrinth seal in this location with a 0.005 in. typical radial

clearance has an estimated leakage flow factor of 0.007.

The long term seal life can not be predicted accurately due

to the limited endurance testing of 50 hr. However, based

on the excellent condition of the test seal and rotor after

50 hr of testing, it is anticipated that the seals should easily

meet the JTAGG II engine test requirement.

Introduction

The JTAGG II engine is an advanced gas generator,

funded by the U.S. Army, Navy and Air Force, being

developed by AlliedSignal Engines in Phoenix. This

engine uses two identical brush seals, in tandem, located

aft of the high pressure compressor (HPC) as shown in

Fig. 1. The two brush seals separate the hot HPC back face

cavity from the relatively cooler buffer air for the number 2

forward carbon ring seal. The IRP operating conditions

for this seal in the JTAGG II engine are estimated to be

50 000 rpm (899 ft/sec) speed, 175 psid air to air pressure

differential and 1200 °F air temperature. A purchase order

was issued to NASA Lewis Research Center to test these

two brush seals in a seal rig. These seals were tested

between March, 1996 and January, 1997. Because the

NASA seal rig did not have the capability to test the

This paper is declared a work of the U.S. Government and is not

subject to copyright protection in the United States.

seal to maximum engine speed, pressure differential and

temperature simultaneously, testing was conducted up to

the maximum operating conditions shown in Table I.

Following a description of the test apparatus, the

procedures, air leakage and wear results will be discussed

for the static, performance, brush seal offset, rotor run out,

and endurance tests.

Brush Seal and Rotor Description

A total of 4 sets of brush seals, each set consisting of

two seals, and four rotors were designed and procured for

rig testing. A typical test brush seal and its nomenclature

is shown in Fig. 2. Two brush seal configurations,

conventional low hysteresis and advanced low hysteresis

design, were used in testing. The advanced low hysteresis

design brush seals were used only in the endurance tes t ,

which included pre- and post-endurance performance

testing. The seal was designed to have a clearance with the

rotor at build and a line to line to slight interference at

engine IRP speed of 50 000 rpm. Brush seal bristle and

side plate material were Haynes 214 and INCO 625,

respectively. The test rotor shown in Fig. 3 was made of

INCO 718 and coated with EG&G Sealol's proprietary

Triboglide coating. Triboglide is essentially a chromium

carbide coating with barium calcium fluoride, high

temperature solid lubricant, dispersed in the coating. The

rotor o.d. is 4.121 in. One rotor had a 0.003 in. radial run

out in its o.d. relative to the i.d. for the rotor run out test.

All rotors were balanced at low speed and then trim

balanced in the test rig prior to testing with a seal.

Seal Rig Description

The NASA Lewis seal rig, shown in Fig. 4, was used

for testing the brush seals. An adapter was made to fit in

the existing seal holder to mount the smaller diameter

brush seals for the JTAGG II engine. In addition, a new

American Institute of Aeronautics and Astronautics
plenum was made to direct the flow to the o.d. of the rotor

which was also smaller. An enlarged view of the seal and

rotor installation in the NASA Lewis rig is shown in Fig. 5.

For the seal offset test, a second adapter was made with a

radial offset of 0.003 in. between the adapter o.d. and i.d.

Static Air Leakage Test

This test characterized the brush seal air leakage as a

function of air to air pre ssure differential and air temperature

under static conditions. Conventional low hysteresis brush

seals S/N 1 and 2 were used in this test as the low pressure

side and high pressure side seals, respectively.

Leakage Performance

Figure 6 shows the air leakage flow factor as a

function of pressure drop across the seal at air temperatures

of 70, 500, 800 and 1100 °F. The flow factor, d_,is defined

as follows:

_=lil× _(Tavg +459.67) lbm-in.--x_-

(Pu ×Di) ' lbf-s

where

rn = Air leakage flow rate, lbm/sec

Tavg = Average air temperature upstream of

the seal pair, °F

Pu = Air pressure upstream of the seal pair,

psia

D i = Outsidediameterofthesealrotor,4.121 in.

As shown in Fig. 6, the flow factor decreases as air

temperature increases. This is due to a thermal mismatch

between the seal and tlie rotor, which reduces the

clearance at higher temperature. Also at higher temperature,

viscosity of the air increases, which offers higher resistance

to air flow and therefore a further reduction in air flow

rate through the seal. Figure 6 also shows that the flow

seems to start choking at about 50 psid (a pressure ratio of

about 3.4).

Performance Test

This test characterized the brush seal air leakage as a

function of speed, air to air pressure differential and air

temperature. Conventional low hysteresis brush seals

S/N 1 and 2 were used in this test as the low pressure side

and high pressure side seals, respectively.

Leakage Performance

Figures 7 to 10 show plots of air leakage flow factor

as a function of pressure ratio across the seal pair for speeds

from 0 to 45 000 rpm at air temperatures of 70, 500, 800,

and 1100 °F. The pressure ratio, Pr, is defined as follows:

Pr = Pu/Pd ,

where

Pu = Air pressure upstream of seal pair, psia

Pd = Air pressure downstream of seal pair, psia

The results of the speed ramp up and down cycle

test are plotted in Fig. 11. The following observations

can be made from these plots:

• The flow factor decreases as speed increases. As

the speed increases, the rotor grows due to centrifugal

force, therefore reducing the clearance between the seal

and the rotor. Lower clearance leads to a decrease in air

flow rate.

• As in the static test, the flow factor decreases as

temperature increases.

• The air leakage flow begins to choke at a pressure

ratio between 3 and 4.

• For the maximum test conditions of 45 000 rpm

speed, 120 psid pressure differential and 800 °F air

temperature, a flow factor of less than 0.002 was

measured, which is below the flow factor goal of 0.004.

• It is feasible to extrapolate the flow factor, with

reasonable accuracy, for the engine IRP operating

conditions of 50 000 rpm speed, 175 psid pressure

differential and 1200 °F air temperature from the results

plotted in Figs. 9 and 10. Because the flow is fully choked

before the maximum rig operating condition of 45 000 rpm,

120 psid and 800 °F is reached and because the engine IRP

operating condition is higher than the rig maximum

operating condition, it is reasonable to assume that the

flow is also choked at the engine IRP operating condition.

Therefore the flow factor at the engine IRP operating

condition is expected to be similar to that at the rig

maximum operating condition. It is estimated that the

engine air leakage flow factor at the IRP operating

conditions will be less than 0.002 for a small rotor run out

of 0.0005 in., similar to that in the rig. For higher rotor run

out, the flow factor is expected to increase in value. This

is explored later in the rotor run out test. The leakage flow

factor for the brush seal is significantly less than the

leakage flow factor of 0.007 for the labyrinth seal of

0.005 in. radial clearance at the JTAGG II operating

conditions.

• The seal pair (S/N 1 and 2) tested, showed

considerable hysteresis as is evident in Fig. 11. This test

J

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American Institute of Aeronautics and Asla'onantics
wasrunataconstant pressuredifferential acrosstheseal

of 60psidandaconstant temperatureof 800°F,while

speedwasrampedupandthendown.Theairleakageflow

wasmeasuredasafunctionofspeed. Theplotshowsthat

theflowfactorwashigherduringtherampdowncycle as

compared to the ramp up cycle, It is anticipated that on the

ramp up cycle bristles move out due to the rotor centrifugal

growth, thermal mismatch between the seal and the rotor,

rotor run out and seal offset, etc. It appears that during the

ramp down cycle, the seal bristles do not retum to their

original position, resulting in higher leakage rates. This

would imply that the frictional force between the low

pressure side plate and the bristles is greater than the

restoring force in the bristles.

This problem was discussed with the seal vendor

EG&G Sealol. A new set of seals (S/N 5 and 7), which

incorporated their latest technology, was procured and

tested later in the program (before the endurance test).

These seals had two features: (1) the low pressure side

plate was relieved where it contacts the bristles (similar to

the conventional low hysteresis design), and (2) a thin

deflector plate was inserted between the high pressure side

plate and the bristles. Elements of this design are covered

under U.S. patent number 5401036, but the details of some

additional features proprietary to EG & G Sealol are not

available. These features have shown reduction in seal

hysteresis in rig testing at EG&G Sealol. 1Figure 2 shows

a conceptual view of the advanced test seal. A photograph

of the advanced low hysteresis seal is shown in Fig. 12.

The pre-endurance speed ramp up and down test

results from the advanced S/N 5 and 7 seals are shown in

Fig. 13. The difference between the flow factor at speed

ramp down and ramp up is much smaller as compared to

the conventional seals, S/N 1 and 2 (refer to Fig. 11).

The advanced seal pair was subjected to speed ramp

up and down test after completion of 50 hr endurance test.

The results of the test are shown in Fig. 14. The seals

showed evidence of hysteresis in this test similar to the

conventional design (Fig. 11).

These results were discussed with the seal vendor

EG&G Sealol. At this point, an acceptable explanation for

this phenomenon has been not determined. However, as

per EG&G Sealol, the advantage of advanced seal design

lies in its ability to maintain a low contact load between the

bristle pack and the rotor during high deflection transients

like traversing through critical speeds and maneuver loads.

This results in a reduction in bristle wear, therefore an

increase in seal life.

Seal and Rotor Wear

The total run time accumulated during performance

testing was 10.4 hr.

The post test inner diameters for seals S/N 1 and 2,

were slightly larger than the pretest inner diameters by

0.0024 and 0.0002 in., respectively. The inner diameter

change could be the result of wear, inward flowering of

bristles during performance testing, and measurement

errors. An optical comparator was used to measure the

brush i.d at eight circumferential locations. Stray bristles,

inaccuracy in locating the center of the brush, and the

compliant nature of the bristles brings uncertainty to the

brush i.d. measurement. This subject will require further

investigation to reach a final conclusion. In spite of these

uncertainties, leakage flow factor goals were met, as

described earlier.

The rotor had two visible wear tracks on its mating

o.d. corresponding to the two seals tested in tandem. The

wear track width of 0.1181 in. was much larger than the

0.050 in. approximate bristle pack width at build. This

may be a result of axial bristle movement during

performance testing and the relative axial thermal

movement between the bristles and the rotor. A

profilometer was used to measure the rotor wear at eight

approximately equally spaced circumferential locations.

The wear track depth varied from 0.00021 to 0.00078 in.

for track 1 and 0.00051 to 0.00119 in. for track 2. Tracks

1 and 2 correspond to the seal on the downstream side (low

pressure side, seal S/N 1) and upstream side (high pressure

side, seal S/N 2), respectively. The coating wear does pose

some concern. This was discussed with the coating experts

at EG&G Sealol, whose proprietary Triboglide coating

was used on the rotor. Triboglide, as mentioned earlier, is

achromium carbide coating with barium calcium fluoride,

high temperature solid lubricant, dispersed in the coating.

A further investigation revealed that the coating vendor

had used a higher percentage of the solid lubricant (about

11 percent) than Sealol's recommendation (6 percent).

Also the surface finish of the rotor at build (46 ktin.) was

much higher than the print requirement (8 [xin.). It is felt

that improvement in wear coating and surface finish will

help in reducing coating wear. 2

Energy Dispersive Spectroscopy (EDS) was used to

look for material transfer from the bristles to the rotor and

vice versa. In the Triboglide coating the following elements

can be identified: chromium, nickel, calcium and barium.

The other constituents of Triboglide, carbon and fuorine,

have atomic numbers that are too low to be detected on the

Princeton Gama-Tech System 4 that was used. In the areas

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American Institute of Aeronautics and Astronautics
observed, there was no significant difference in the amount

of the elements found inside or outside the tracks on the

rotor, with the exception that a trace amount of Aluminum

was found in the track. Aluminum is a constituent of the

bristle material, Haynes 214. Note that iron, aluminum

and yitrium are the elements of Haynes 214 that are not

common with Triboglide. Conversely, the elements in

Triboglide that are not common with Haynes 214 are

calcium, barium and fluorine. In the areas observed on

bristle pack of S/N 1 and 2 seals, no calcium, barium or

fluorine was found. However, since there are like elements

in the bristles and rotor coating (nickel and chromium) no

conclusion can be drawn stating whether the Triboglide

transferred to the bristle tips or not. It does not appear that

bristle material transferred to the rotor. Using the

backscatter mode of the Scanning Electron Microscope

(SEM), it was observed that there were less black spots in

the track. In this mode on the SEM, lower atomic elements

are darker and it is likely that some calcium barium

fluoride has been worn from the rotor. Also through the

SEM, smearing type wear was observed on the bristle tips.

Brush Seal Offset Test

This test simulates the maneuver type condition in a

flight engine. In this test, each seal of the seal pair was

radially offset by 0.003 in.with respect to the rotor. This

was accomplished by machining a 0.003 in. radial offset

in the seal adapter i.d. relative to the adapter o.d., which

mates with the seal holder as shown in Fig. 3. Conventional

low hysteresis brush seals S/N 3 and 4 were used as thelow

pressure side and high pressure side seals, respectively.

Data was taken to characterize seal air leakage as a

function of time. The operating conditions were 45 000 rpm

speed, 60 psid air to air pressure differential and 1100 °F

air temperature for the first 10 hr, and 45 000 rpm,

120 psid and 800 °F for the final 10 hr of running. The

seals and rotor were removed from the rig and inspected

after the first 10 hr and again after the test was completed.

Leakage Performance

Figure 15 shows a plot of air leakage flow factor as a

function of accumulated run time. The steps or breaks in

the data that occur every 5 hr are indicative of taking the

data over 4 run days of 5 hr each. A slight increase in air

leakage was observed with time, indicating that the

clearance between the seal bristles and rotor did not

change significantly in 20 hr of running. The maximum

flow factor measured was 0.0023, which is below the

design goal of 0.004.

Seal and Rotor Wear

The total run time accumulated during offset testing

was 20 hr.

After 10 hr of run time, the bristle i.d. for seals S/N 3

and 4 was larger than at pretest by 0.0029 and 0.0012 in,.

respectively. After 20 hr of run time, the bristle i.d. for seal

S/N 3 was larger than at pretest by 0.0006 in. and for seal

S/N 4 seal was smaller by 0.0014 in. A po ssible explanation

for this could be that bristles flowering inwards

compensated for wear on the bristles as the run time got

closer to 20 hr. However, as mentioned earlier, the brush

seal i.d. measurement can not be accurately made. In spite

of these uncertainties, leakage flow factor goals were met,

as described earlier.

The wear track width of 0.15 in. was much larger than

the 0.050 in. approximate bristle pack width at build. This

may be due to axial bristle movement during offset testing

and the relative thermal axial movement between the

bristles and the rotor.

The wear track depth, after 10 hr of run time, varies

from 0.00009 to 0.0010 in. for track 1 and 0.00005 to

0.00017 in. for track 2. The wear track depth, after 20 hr

of run time, varies from 0.0002 to 0.001 in. for track 1 and

0.00001 to 0.00018 in. for track 2. Rotor wear in the seal

offset test was less than the rotor wear in the performance

test, which may be due to the finer surface finish (8 ktin.)

of the rotor used in the seal offset test. The rotor wear

seems acceptable, since the seal leakage flow factor did

not change significantly after 20 hr of run time.

Rotor Run Out Test

This test simulated the transient conditions in a flight

engine, such as traversing through critical speeds and fast

accelerations and decelerations, etc. The rotor o.d, had a

radial run out of 0.003 in. with respect to its i.d. pilot.

Conventional low hysteresis brush seals S/N 6 and 8 were

used in this test as the low pressure side and high pressure

side seals, respectively. Data was taken to characterize the

seal air leakage as a function of run time. During the first

10 hr the operating conditions were 40 000 rpm speed,

60 psid air to air pressure differential and 1100 °F air

temperature. Then the seals and rotor were removed from

the rig and inspected. After reinstallation, testing continued

at operating conditions of 40 000 rpm, 95 to 100 psid

across the seal, and 800 °F inlet air temperature for another

10 hr. The seal and rotor were removed from the rig and

given a final inspection. Although trim balanced,

40 000 rpm was the maximum shaft speed obtainable with

the 0.003 in. radial run out in the rotor, due to rig vibration

limits.

Leakage Performance

Figure 16 shows a plot of air leakage flow factor as a

function of time. The leakage flow factor approximately

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American Institute of Aeronautics and Astronautics
doubled, fromabout0.002to0.004,after5hrofrunning.

Inthenext15hrofrunning,leakageflowfactorincreased

slightlytoalittleover0.004.Itcanbeconcluded that after

5 hr of running, the flow factor stabilized. These results

also indicate that rotor run out had a much more pronounced

effect on air leakage as compared to seal offset. Again, the

steps in the data reflect the four 5 hr run days.

Seal and Rotor Wear

The total run time accumulated during rotor run out

testing was 20 hr.

After 10 hr of run time, the bristle i.d. for seals S/N 6

and 8 was larger than at pretest by 0.0082 and 0.0027 in.,

respectively. After 20 hr of run time, the bristle i.d. for

seals S/N 6 and 8 was larger than at pretest i.d. by 0.0064

and 0.0012 in., respectively. A possible explanation for

this could be that the bristle flowering inwards was larger

than the bristle wear as run time approached 20 hr. In spite

of the uncertainty in the brush seal i.d. measurement, the

increase in the brush seal i.d. for the rotor run out test was

significantly more than for previous tests.

The rotor, which had an initial surface finish of 21 to

23 _tin., showed wear as well as material build up in its

wear tracks. The maximum groove depth in wear tracks 1

and 2 after 10 hr of run time was 0.00052 and 0.00087 in.,

respectively. The maximum build up in wear tracks

numbers 1 and 2 after 10 hr of run time was 0.00068 and

0.00059 in., respectively. The maximum groove depth in

wear tracks 1 and 2 after 20 hr of run time was 0.00092 and

0.00088 in., respectively. The maximum build up in wear

tracks 1 and 2, after 20 hr of run time, was 0.00055 and

0.0005 in., respectively. The maximum groove depth

occurred at the same location as the maximum rotor run

out. The maximum build up occurred at 90 ° from the

maximum groove depth. The build up on the rotor is

indicative of metal transfer.

EDS plots for the rotor after twenty hours run time

showed the presence of aluminum in both the low and high

pressure side wear tracks on the rotor. Aluminum is an

element found in Haynes 214 bristle material, but not

found in Triboglide coating on the rotor. Based on this

evidence and build up of material seen on the rotor wear

tracks, it can be concluded that Haynes 214 bristle material

transferred to the rotor.

The effect of material transfer on seal performance is

difficult to assess. However, as mentioned earlier, leakage

flow factor increased from -0.002 to 0.004 after 5 hr of run

time and then remained close to the leakage flow factor

goal of 0.004 for the remaining 15 hr of the test. Though

the testing was limited to 20 hr, the results indicate that

brush seals have good potential for long life applications.

Endurance Test

A 50 hr endurance test was conducted to evaluate the

long term air leakage characteristics of the brush seal.

Advanced low hysteresis brush seals S/N 5 and 7 were

used in this test as the low pressure side and high pressure

side seals, respectively. This test was conducted at operating

conditions of 45 000 rpm speed, 60 psid air to air pressure

differential and 1100 °F air temperature for the first 25 hr

and 45 000 rpm, 90 to 103 psid and 800 °F for the final

25 hr of running. The seals and rotor were inspected after

the first 25 hr of testing and again at test completion.

Leakage Performance

Figure 17 shows a plot of air leakage flow factor as a

function of time. There was no significant change in flow

factor after 50 hr of run time. The flow factor was close to

0.002 for the entire duration of the endurance test. This is

well below the flow factor goal of 0.004.

Seal and Rotor Wear

The total run time accumulated during endurance

testing was 50 hr.

After 25 hr of run time, the bristle i.d. for seal S/N 5

seal was larger by 0.0024 and for seal S/N 7 was smaller

by 0.00145 in. After 50 hr of run time, the bristle i.d.

for seals S/N 5 and 7 was smaller than at pretest by 0.0051

and 0.0052 in., respectively. A possible explanation for

this could be that bristle flowering is larger than bristle

wear. However, as mentioned earlier, the brush seal i.d.

measurement can not be made accurately. It was fairly

consistently observed that the standard deviation of the

brush seal i.d. measurements increased as run time was

accumulated.

Visual inspection showed that several rows of bristles

protrude radially inwards more than their neighboring

bristles over two 30 to 40 ° sectors on brush seal S/N 5 and

over an -90 ° sector on brush seal S/N 7. Therefore change

in brush i.d. is not a good indicator of bristle wear, but in

this case indicates inward flowering of the bristles.

Since the i.d. measurement was not a good indicator

of bristle wear, a method developed by Fellenstein 3 was

adopted. In this method the low pressure side plate of each

brush seal was marked in four equally spaced circumferential locations. Prior to endurance testing, calibrated

magnified photographs were taken of the bristle length

extending beyond the low pressure side plate and of the

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American Institute of Aeronautics and Astronautics
bristlelengthextendingbeyondthedeflectorplateon the

high pressure side at each of the marked locations.

Photographs were taken again after testing. Pretest and

post-test bristle angle and length were measured from the

photographs to determine bristle wear. Visual averaging

of bristle lengths was used where necessary. Data from the

sector where bristles were extending in radially were

excluded from the average of bristle wear. The average

bristle wear was 0.023 and 0.016 in. on the low pressure

side and 0.013 and 0.009 in. on the high pressure side of

brush S/N 5 and 7, respectively. The bristle angle change

was 2.3 ° for seal S/N 5 and 1.5 ° for seal S/N 7. Since

measurements indicate that flowering compensates for

the considerable bristle wear over the 50 hr test time, it is

concluded that the clearance did not change significantly.

This is also evidenced by the small change in leakage flow

factor over the test duration.

The rotor, which had an initial surface finish of

24.3 pin., showed no grooves, but only material build up

in its wear tracks. The maximum build up in wear tracks

1 and 2 after 25 hr of run time was 0.00152 and 0.00041 in.,

respectively. The maximum build up in wear tracks 1

and 2 after 50 hr of run time was 0.00078 and 0.00031 in.,

respectively. The maximum buildup of material on the

tracks has decreased from the buildup value after the first

25 hr of run time. The amount of buildup is small and does

not have any significant impact on the seal performance,

as is evident from the relatively constant flow factor of

-0.002 for the entire duration of the endurance test.

Aluminum, an element in the Haynes 214 bristle

material, but not in the Triboglide coating, was found in

both rotor wear tracks after 25 and 50 hr of testing through

EDS analysis. Basedonthis evidence andtheprofllometer

traces which showed material build up in rotor wear tracks,

it can be concluded that the Haynes 214 bristle material

transferred to the rotor. Also, some hairline cracks were

found in the coating material on the low pressure side of

the low pressure track; the cracks were parallel and in the

axial direction. EDS was also done on one brush seal,

S/N 5, to examine the bristle tips. Some debris was found

which contained calcium, which means the debris may be

Triboglide. However, there are other elements in the debris

that may indicate it is just dirt picked up from handling.

Conclusions

1. Based on the extrapolated test results, the brush

seal design should be able to meet the air leakage flow

factor goal of less than 0.004 for the engine operating

conditions. This is a significant improvement over the

leakage flow factor of 0.007 for a labyrinth seal with a

0.005 in. radial clearance.

2. The excellent condition of and low wear observed

on the post test seals and rotors indicate that the brush seals

have an excellent potential for long life applications.

3. Some material transfer occurred between the seal

bristles and the rotor coating; however, the material transfer

was small.

4. Both the conventional and advanced seal designs

showed considerable hysteresis. Further research is

recommended to reduce seal hysteresis. This will help

further reduce the seal air leakage flow rate.

5. Rotor and seal wear and leakage performance

were more effected by rotor run out than by seal offset.

6. The leakage flow factor decreases as speed and

temperature increase due to a reduction in seal clearance

and increased air viscosity.

7. Air leakage flow began to choke at pressure ratios

between 3 and 4.

8. In the endurance test, the bristle wear was

compensated by inward flowering of the bristles, resulting

in minimal clearance change and, therefore, only a small

change in leakage flow factor over 50 hr of testing.

9. Both brush i.d. and bristle wear measurements

should be made to assess bristle flowering and wear in

brush seals.

Acknowledgments

The work described in this paper reflects efforts

supported by the Tri-services JTAGG program of the U.S.

Army, Aviation Applied Technology Directorate(AATD),

Naval Air System Command Propulsion and Power

Engineering Dept. (AIR4.4), and Wright Laboratories

Aero-propulsion and Power Directorate (WL/PO). The

AATD program manager is S andra Hoff and the mechanical

components technical monitor is Phil LaFarriere for the

JTAGG II program.

The authors acknowledge the contributions of NASA

Lewis Research Center, Cleveland, OH, where all the

testing was conducted and the operations support of Joseph

Flowers of the Army Research Laboratory, Vehicle

Technology Center-Lewis Site. The authors also thank

EG&G Sealol, Engineered Products Division,

Warwick, RI, where all the test seals were fabricated.

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American Institute of Aeronautics and Astronautics
References

1. Short, J.F., Basu, P., Datta, A., Loewenthal, R.G.,

Prior, R.J., "Advanced Brush Seal Development,"

AIAA Paper No. 96-2907, 1996.

2. Ferguson, J.G., "Brushes as High Performance Gas

Turbine Seals," ASME Paper 88-GT-182, 1988.

3. Fellenstein, J.A., "A New Tribological Test for

Candidate Brush Seal Materials Evaluation," NASA

TM-106753, 1995.

TABLE I.--NASA RIG TEST AND JTAGG II ENGINE OPERATING CONDITIONS

Operating condition NASA rig NASA rig NASA rig JTAGG H

engine

Maximum speed, rpm (ft/sec) 45 000 (809) 45 000 (809) 45 000 (809) 50 000

(899)

Maximum air to air pressure 60 120 145 175

differential, psid

Maximum air temperature, °F 1100 800 70 1200

Sump Seal

High-Pressure Upstream Downstream Buffer Air

Compressor Brush Seal Brush Seal

Backface Cavity Annulus

No. 2 /

Forward

Brush Seal

Pair

High-Pressure

Compressor

No. 2 No. 2

Air Vent Forward Roller

Cavity Carbon Ring Bearing

Seal

G7999-1

Figure 1 ._JTAGG II No. 2 forward brush seal arrangement.

-E

7

American Institute of Aeronautics and Astronautics
Relief

annulus

Conventional Advanced

low-hysteresis design low-hysteresis design

(S/N 1,2, 3, 4, 6, and 8) (S/N 5 and 7)

1. High-Pressure Side Plate, INCO 625

2. Bristle Pack, Haynes 214,

Bristle Diameter 0.0028 Inch

3. Low-Pressure Side Plate, INCO 625

4. Deflector Plate, INCO 625

Figure 2._JTAGG II Typical test brush seal and its

nomenclature.

Triboglide

coating

Rotor material: INCO 718

4.121

in.

Balance

ring

Balance

ring

Tapped holes for

trim balancing

the rotor

0.660 in.

Figure 3.---JTAGG II Brush seal rig test rotor.

8

American Institute of Aeronautics and Astronautics
High-pressure

air supply

I

Figure 4.--NASA Lewis seal rig cross section.

/-- Brush seal

holder

Pressure

chamber

(high-pressure

side)

Seal

rotor

Clamping

nut

r Plenum

_-- Cover

,/ plate

9

American Institute of Aeronautics and Astronautics
Seal adapter Clamp

plate

Brush seal

Optional

seal

installation

Figure 5.--NASA Lewis seal rig brush seal and rotor installation for JTAGG II.

10

American Institute of Aeronautics and Astronautics
0

o

O)

O_

..J

0.008

0.006

0.004

0.002

0.000

i r i i

Inlet air temperature,

oF

o 70

[] 500

800

1100

0

o

I i _ i i I J i i

I I I I I I q I I I

50 1O0

Pressure drop across seal, psid

Figure 6.--Static test: leakage flow factor vs. pressure drop across seal as a function of inlet air

temperature. Conventional low hysteresis brush seals, SIN 1 and 2.

150

0.006

+a

0.004

_o

I1)

0.002

Shaft speed,

rpm

o 0

[] 25 000

40 000

45 000

0 0

0

[] []

0 []

[]

0

[]

0.000 _ I I I ' I I I I I I I I I ]

0 2 4 6

Average pressure ratio

Figure 7.--Performance test: average flow factor vs. average pressure ratio as a function of

speed at 70 °F inlet air temperature. Conventional low hysteresis brush seals, S/N 1 and 2.

8

11

American Institute of Aeronautics and Astronautics
O

O

O3

0.006

0.004

0.002

Shaft speed,

rpm

o 0

[] 25 000

45 000

[]

O

0.000 , , , I , , , t , , , I , , ,

0 2 4 6

Average pressure ratio

Figure 8.--Performance test: average flow factor vs. average pressure ratio as a function of

speed at 500 °F inlet air temperature. Conventional low hysteresis brush seals, SIN 1 and 2.

0.006

0.004

_o

O3

0.002

Shaft speed,

rpm

o 0

[] 25 000

o 45 000

O

O

0 0 [] [] [3

0

%

%

0.000 _ , , I , , _ I , , , I , , ,

0 2 4 6

Average pressure ratio

Figure 9.--Performance test: average flow factor vs. average pressure ratio as a function of

speed at 800 °F inlet air temperature. Conventional low hysteresis brush seals, SIN 1 and 2.

8

0.006

O

iiO

O3

G)

0.004

0.002

Shaft speed,

rpm

o0

40 000

o 45 000

0

0

0

A A A 0

0 0

0.000 i t I J I, , I , , , I , , ,

0 2 4 6 8

Average pressure ratio

Figure 10.--Performance test: average flow factor vs. average pressure ratio as a function of speed

at 1100 °F inlet air temperature. Conventional low hysteresis brush seals, SIN 1 and 2.

12

American Institute of Aeronautics and Astronautics
0.006

0.004

o

_ 0.002

6 A A

A

A

• A

o Increasing speed

A Decreasing speed

0.000 , , _ _ , , , , , I , , , , , , , , , I , , , , , _ _ _ , I , , , _ , , , , _ I , , , , _ , _ , ,

0 10 20 30 40

Runner speed, Krpm

Figure 11.--Effect of speed ramp up and down on leakage flow factor on conventional low

hysteresis brush seals, S/N 1 and 2. Inlet air temperature, 800 °F; pressure drop across seals,

60 psid.

5O

Figure 12.--Advanced low hysteresis brush seal.

13

American Institute of Aeronautics and Astronautics
0.006

0.004

0

_ 0.oo2

-1

# @

o Increasing speed

Decreasing speed

$

$

0,000 , i , i i .... L ......... I , , i , = i i , _ I , i i i = i i i t I i , _ , , I J , ,

0 10 20 30 40 50

Runner speed, Krpm

Figure 13.mEffect of speed ramp_p and down on leakage flow factor for advanced low hysteresis

brush seals, S/N 5 and 7, prior to endurance test. Inlet air temperature, 800 °F; pressure drop

across seals, 60 psid.

0.006

0.004

_o

__ 0.002

.J

o Increasing speed

_ Decreasing speed A

O

o 8 o

o o

o o

0.000 t , , , , , , , , I ......... I ......... I .... , _ , , , i .........

0 10 20 30 40 50

Runner speed, Krpm

Figure 14.--Effect of speed ramp up and down on leakage flow factor for advanced low hysteresis

brush seals, SIN 5 and 7, after endurance test. Inlet air temperature, 800 °F; pressure drop across

seals, 60 psid.

0.0030

0.0020

0.0010

ooo o

ooo

oo

o

O00aOOOOOOO4BO OO000000000000000000000000_0

0

0 0 0

0 0 CO00000000

0000 0 0 0

0 0

o

0.0000 I I I I

0 4 8 12 16 20 24

Accumulated time, hr

Figure 15._Seal offset test: flow factor vs. accumulated run time. Conventional low hysteresis

brush seals, S/N 3 and 4.

14

American Institute of Aeronautics and Astronautics
0.0050

0.0040

0.0030

O

0.0020

0.0010

0.0000

o

o

o °

o

00(3 ooo°ooooo OOOOOO

o

o oo

o oooooo oooooooo oo

O OOOOOoooOooo OOC_ O

ooo O ° o o o

oo o

o ° o

oo

o ° o

o

I t t I I

0 4 8 12 16 20

Accumulated time, hr

Figure 16.--Rotor runout test: flow factor vs. accumulated run time. Conventional low

hysteresis brush seals, SIN 6 and 8.

24

0.0030

0.0020

0.0010

%=====o

® o (3o _0¢_ml_o o ° °o

0lifO o0 Olll)_O_ o oooo _ OCXJDOODOGIX_O

• lO O ¢_DO OODO _19 OOO OQO

(I_ O0 _ O0 0

o oo_ °_ o o oo__

g

o

0.000 E , , , , , E , I i

0 5 10 15 20 25 30 35 40 45 50 55

Accumulated time, hr

Figure 17.--Endurance test: flow factor vs. accumulated run time. Advanced low hysteresis

brush seals, S/N 5 and 7.

15

American Institute of Aeronautics and Astronautics
Form Approved

REPORT DOCUMENTATION PAGE OMB No. 0704-0188

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,

gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this

collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson

Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE

July 1997

4. TITLE AND SUBTITLE

JTAGG II Brush Seal Test Results

6. AUTHOR(S)

Gul K. Arora and Margaret P. Proctor

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Lewis Research Center

Cleveland, Ohio 44135-3191

and

U.S. Army Research Laboratory

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

and

U.S. Army Research Laboratory

Adelphi, Maryland 20783-1145

3. REPORT TYPE AND DATES COVERED

Technical Memorandum

5. FUNDING NUMBERS

WU-242-70_J

1L161102AH45 or

1L162211A47A

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-10725

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM-107448

AIAA-97-2632

ARL-TR-1397

11. SUPPLEMENTARY NOTES

Prepared for the 33rd Joint Propulsion Conference and Exhibit cosponsored by AIAA, ASME, SAE, and ASEE, Seattle, Washington,

July 6-9, 1997. Gul K. Arora, Allied Signal Engines, 111 South 34th Street, Phoenix, Arizona 85072-2181 and Margaret P. Proctor,

NASA Lewis Research Center. Responsible person, Margaret P. Proctor, organization code 5950, (216) 977-7526.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Categories 07 and 20

This publication is available from the NASA Center for AeroSpace Information, (301) 621-0390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

The Tri-services JTAGG II engine uses two identical brush seals, in tandem, located aft of the high pressure compressor.

The engine operating conditions, at intermediate rated power (IRP), for this seal are estimated to be 50 000 rpm (899 ft/sec_

speed, 175 psid air to air pressure differential and 1200 °F air temperature. The testing was comprised of static air leakage,

performance, seal offset, rotor run out tests and a 50 hr endurance test in the NASA Lewis seal rig. Based on the test

results, it is concluded that the brush seal design should be able to meet the air leakage flow factor goal of less than 0.004

for the engine IRP operating conditions. For the 4.12 in. i.d. labyrinth seal, 0.005 in. typical radial clearance, at the JTAGG

II operating conditions, the leakage flow factor is 0.007. The long term seal life can not be predicted accurately due to the

limited endurance testing of 50 hr. However, based on the excellent condition of the test seal and rotor after 50 hr of

testing, it is anticipated that the seals should easily meet the JTAGG II engine test requirement.

14. SUBJECT TERMS

Brush seals; Seals; High temperature; Air

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

18. SECURITY CLASSIFICATION

OF THIS PAGE

Unclassified

19. SECURITY CLASSIFICATION

OF ABSTRACT

Unclassified

15. NUMBER OF PAGES

17

16. PRICE CODE

A03

20. LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102