AIAA-91-2032 Aerodynamic Losses due to Pressure Side Coolant Ejection in a Transonic Turbine Cascade H. L. Moses, T. Kiss and R. Bertsch Virginia Polytechnic Inst. and State Univ. Blacksburg, VA B. A. Gregory General Electric Aircraft Engines, Cincinnati, Ohio

27th AIAA/SAE/ASME Joint Propulsion Conference

June 24-26, 1991 -- Sacramento, CA

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AIM-91-2032-CP

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AERODYNAMIC LOSSES DUE TO PRESSURE SIDE COOLANT EJECTION IN A TRANSONIC TURBINE CASCADE

Hal L. Moses' Tibor Kiss"

Remi Bertscht Virginia Polytechnic Institute and State University, Blacksburg, Virginia

Brent A. Gregory$ General Electric Aircraft Engines, Cincinnati , Ohio

A b s t r a c t Experiments were conducted in a blow-down-type

wind tunnel on two-dimensional cascades of cooled, transonic turbine blades, in the exit Mach number range of 0.7 through 1.4 and blowing rate range of 0 through 1.7. The coolant, which was COz to obtain the proper density ratio with cold flow, entered the blades through holes on the spanwise ends, and exited tan- gentially through slots on the pressure side, near the trailing edge. Only aerodynamic aspects of the flow were studied. The measured data, which was reduced to the mass averaged total pressure loss as a function of isentropic Mach number,show only a slight effect of the coolant flow rate. Generally, the effect of cooling on the losses was within the experimental uncertainty.

Nomenc la tu re

P P T U M Me,"", B

Y

w

2

C

1 pb p d m A

pressure density temperature velocity Mach number time averaged exit Mach number blowing rate axial coordinate in the test section transverse coordinate in the test section axial chord pitch blowing pressure pressure in COz distributor mass flow rate per blade passage mass averaged total pressure loss coefficient

Subscr iDts t C coolant U upstream of the cascade e

stagnation value of a variable

exit values of flow variables

'Professor, Member AIAA **Graduate Student, Member AIAA tGraduate Student $Manager - Turbine Aero & Cooling Technologies Cwwght @ 1991 Ameman lnstltule 01 Aeronautics and

Aslroneulics. Inc. All rights reaeNed 1

1 2

at the place of traversing behind the probe bow shock

Introduction With the continuing effort to build lighter, highly

efficient jet engines, there is a need to understand and improve the performance of all components. For the turbine there is a need for higher inlet temperatures and higher stage work at high efficiency. Higher tem- peratures generally result in large coolant flow rates, at least with present and expected material limitations. Much of the coolant gas is ejected at or near the trailing edge of the blades, where it can effect the base pressure and the trailing shock pattern. Thua, it is important to know how the coolant ejection affects the aerodynamic losses.

Although experimental results from an actual tur- bine would be deskable, tests are very expensive, and it is difficult, if not impossible, to study details of the flow. Thus, researchers have for many years made use of two-dimensional, stationary cascades. With transonic turbine cascades, however, there is often a problem with flow periodicity, due mainly to reflected waves. Espe- cially when a so-called tailboard is used to set the flow angle and the exit Mach number, the reflected shocks may be strong. Not only the periodicity can be dis- torted, but also the measured losses may be consider- ably higher than they should be. In the present work no tailboard was used, and there was evidence that the non-periodicity effects were quite small.

Several investigators have studied the effect of coolant ejection on losses'-6, but the results are still somewhat inconclusive. The lack of a general conclu- sion is due in part to the different ejection geometry used by different investigators. In most of the previ- ous published work, the coolant waa ejected through a necessarily thick trailing edge.

For the present study, the ejection was through slots in the pressure side near the trailing edge, and no film cooling was implemented. The ejected gas was (202, so that without any heating of the mainflow, the coolant - mainflow density ratio could be maintained the same as that in an actual engine.

Exuer imen ta l ADuaratus The Blades

Two similar kinds of blades were tested, and from here on they will be reffered to as the B and U blades, respectively. Both the B and U blades could be the blades of the rotor of a high pressure turbine stage in a jet engine, if two significant changes had not been made for the sake of the cascade testing. The first, o b vious change was that they were two-dimensional. The second change was that the front portion of the blades were modified in such a way that the inlet blade angle became zero degrees, whereby the turning in the duct of the wind tunnel could be kept lower. Both of these changes, however, were believed to have insignificant effects on the conclusions.

For both the B and U blades the the turning an- gle was 67 degrees, the inlet Mach number around 0.2, the axial chord 1.5 inches, and the span 6 inches. The design outlet Mach number was 1.14 for the B blades and 1.20 for the U blades. For the B blades the throat was located very close to the trailing edge, and for the U blades it was slightly upstream. The trailing edges were rounded with a thickness of 0.038 in. for the B blades and 0.030 in. for the U blades.

Fig.1 I n t e r n a l passages of the cooled blades

Not all the blades were manufactured for cooling, only those behind which measurements were performed. The B and U cooled blades were made essentially in the same way, and their internal passages are shown in Fig.1. The COz was supplied to the blades through two holes of a diameter of 5/16 in. drilled in the span- wise ends. The height of the cavity inside these blades

decreased continuously towards the trailing edge, and at 0.61 in. from the trailing edge it reached the width of the exit slots ( 0.02 in.) . There were 40 exit slots that had rectangular cross sections of 0.020x0.090 inch. They were located on the pressure side near the trail- ing edge with axes parallel to the tangent of the mean camber line. The arrangement of these slots is further detailed in Fig.1.

i /

.

Fig.2 The cascade

For the better understanding of the flowfield, blades instrumented with surface pressure taps were manufactured. Unfortunatelly, the cooling of these blades was not possible for practical reasons. For both the B and U blades two instrumented blades were pro- duced with the same arrangement of the pressure taps. One blade had 9 taps on its suction side: the other had 5 taps on its pressure side and 1 tap right on its trailing edge. All the pressure taps were located at midspan.

The Cascade The previously described blades were put in c a s

cades of eleven blades, Fig.2. The pitch for both c a s cades was 1.4674 inches. The frame of a cascade was two plexiglass endwalls on which the spanwise ends of the blades were mounted by a pin and a bolt. Also end pieces were placed between the endwalls below the first and above the last blade. In this way ten similar passages were produced, enough to obtain good peri- odicity. The cooled blades were the 4th, 5th and 6th from the bottom. Holes were drilled in the endwalls to allow the COz tubes to reach the inlets on the blades.

u-

2

n

-

Fig3 Schemat i c of the cascade wind tunnel

The instrumented blades were the 7th and 8th blades from the bottom, except when the effect of cool- ing on the blade surface pressures was studied. In that case the 4th and 6th were the instrumented blades with the cooled 5th blade between them.

For static pressure measurements, two rows of pressure taps were manufactured on one of the end- walls. The first row was 1/4 in. behind the line of the trailing edges. The number of the taps in this row was eleven, and the spacing between them was one fifth of the pitch. They spanned the 4th and 5th blade pas- sages. The second row consisted of only 3 taps, I inch behind the line of the trailing edges. These three taps spanned the 5th passage from the bottom.

- A more detailed description of the blades and cas-

cades can be found in Singer's' and Bertsch's' work.

Wind Tunnel and Coolant Suuph The cascade was placed in the test section of a

blow-down type supersonic wind tunnel6, Fig.3. The working air was pumped through a dryer and a heat exchanger into a storage tank until the pressure inside the tank built up to somewhere between 12 and 18 atm. Before reaching the test section, the air ran through a safety valve, a control valve, a flow straightener and a cross section transition piece. The safety valve was completely open during the tunnel runs. The control valve was used to adjust the total pressure upstream of the blades to a constant value. Since it was a pneu- matic valve, its response to the governing signal was relatively slow, and therefore a special feedback system bad to be designed for its proper operation'. Behind the transition piece a settling chamber and then the test section with the cascade followed. Note, that there was no tailboard implemented to turn the Bow in the test section, but rather free shear layers downstream of the blades. After leaving the test section, the air proceeded to a muffler and exited into the atmosphere.

..,

3

The schematic of the COz supply line is depicted in Fig.4. The coolant was applied to the blades from a low pressure tank through a pressure regulator. The same COz tank pressure was used for every run ( 50 psig ), and the tank was large enough for decreasing less than 10 percent of that pressure during one run. On the duct going to the blades the first element was a solenoid valve to turn the blowing on and off. The valve was followed by the pressure regulator, so that the blowing conditions could be maintained constant for one run. The pressure downstream of the pressure regulator was called the blowing pressure, Pb, which roughly deter- mined the volumetric flow rate. The next element was a float-type rotameter to measure the volumetric flow rate. Then the COz proceeded to a relatively large box which distributed it into n i x tubes. Two tubes were led to the COz inlets on each side of each cooled blade.

Measured Variables - Data Aauisition A single, fixed total pressure probe was used for

taking the total pressure just upstream of the cascade, Pt,, since there the flow was uniform. Another total pressure probe traversed parallel to the line of the trail- ing edges of the blades taking the downstream total pressures, Ptz. Pty was led to a pressure transducer, and Ptz and Pt, were lead to a differential transducer. Simultaneous readings were taken on these two trans- ducers at a frequency of 4Olsec. The probe speed was such that with this frequency 328 readings were per- formed in one blade passage, high enough for good res- olution of the total pressure profile. In fact, 800 read- ings were taken on each channel, which means more than two blade passages were traversed. The travers- ing mechanism was made in such a way that the w- ial distance of the probe from the line of the trailing edges could be adjusted. In most cases traversing took place at one of two locations. When the probe traversed along the line of the first row of wallstatic pressure taps (z = 0.25 in.) it was said to be at station 1. When it

BLOWING PRESSURE

m BLADE 4 m - a m BLADE 6

Fig.4 Schematic of the CO1 ~ ~ p p l y system

traversed along the line of the second row (2 = 1.00 in.), then it was said to be at station 2.

A separate system was used for the static pressure measurements, that is for the wall static pressure and the blade surface static pressure measurements. The static pressures were read only one time in each run, by a 32 channel transducer. The scanning of this trans- ducer was fast enough that the readings on the indi- vidual channels could be considered simultaneous. The reading took place at the beginning of the run, at about the same time when the traversing started. Although Pt, was measured on the other data aquisition line, it also was hooked up to the 32 channel trancducer to ensure a reading which took place at exactly the same time as the static pressure readings.

As was mentioned before, the coolant volumetric flow rate was measured by a rotameter. Since the ro- tameter was calibrated for air at standard pressure, a correction of the readings was necessary. For this rea- son the static pressure in the CO2 distribution box, Pd, was measured.

The atmospheric pressure was recorded also, since it was needed for the data reduction.

Data Reduction The Bow conditions in the cascade were set by the

average isentropic exit Mach number, from hereon sim- ply exit Mach number, Me. At the instant of the static Pressure readings Me was the Mach number obtained from the upstream total pressure, Ptu, and from the average of the static pressures over two complete blade Passages at station 1, P,, using the isentropic relation- ship:

there was only one reading on the static pressure taps, a correlation between Pt, and P, had to be established using data obtained from many runs. This way, P, could be obtained for every reading of Pt, in the course of a run. Then, M, could be calculated and averaged over the time of traversing two blade passages. This average is denoted Me,,,,, and is referred to as average exit Mach number.

A similar function between Pt, and the average of the static pressures at station 2 was established. When the local static pressures had to be used during the data reduction, these functions were evaluated with the instanteneous Ptu, therefore the crosswise variation of the static pressures was not incorporated in the results for the losses.

L/

The mass averaged total pressure loss coefficient, A, is defined as

Pt,m + Pi,m, - & P t d cosady Pt,m + pt,m.

where Pil is the total pressure, PtC is the total pressure of the coolant at the coolant exit, m is the massflow rate of the air through one blade passage, mc is the massEow rate of the coolant in one blade, U is the velocity, y is the vertical coordinate, Q is the outlet angle and I is the pitch. For evaluating A with this precise formula one needs to know the mixture density at the measuring station. However, it turned out that a simpler formula

A = ( 2 )

(3) .fd(Pty - Pti)pUcosady A =

Pt,pU cos ady

gave sufficiantly accurate results ( see uncertainties ).It was much easier to evaluate, and therefore it replaced Eq.(2) in the data reduction. ( Incidentally, this a p proximate formula becomes a precise one when coolant

(I)

Because the control of Pt, was not perfect, Pt, and 7-1

therefore P. changed slightly during one run. Since is not applied. ) v

4

Another approximation was that the possible cross- wise changes in the exit flow angle were neglected, and therefore cos a cancelled out. For determining the value of Ptl it was necessary to check whether or not a bow shock was present in front of the traversing probe. The static pressure, PI, in front of the probe bow shock, if this existed at all, was obtained from the correla- tion with the upstream total pressure. For those points of the traverse for which P,,/Pl was less than 1.893 the flow was considered subsonic, and Prl was equal t o the measured Pt2. For those points of the traverse for wich Ptl/Pl was greater than 1.893 the flow was con- sidered supersonic, and the reading on the traversing Pitot tube, P12, was not equal to Pti. In these cases Ptl was determined by an iterative calculation. The Mach number in front of the bow s h o d was assumed to be Mi, then the static pressure and Mach number behind the shock, Pi and M;, could be calculated via the nor- mal shock relations using the static pressure in front of the shock, PI. Plz was then calculated from Mi and P;. P12 was compared to the measured value, and Mi changed until Ptz and P& differed by a sufficiantly low amount. This way the true Ml was determined, and could be used along with PI to calculate Pt1. M I also could be used to determine the static temperature, 2'1, after the total temperature was assumed to be a con- stant for the whole flowfield ( 285K ). Finally, the Mach number and the temperature determined the velocity, U , and the static pressure and the temperature deter- mined the density, p , through the equation of state for

-

- air.

The blowing rate is defined as

(41

where pe and U. are an average density and velocity of the mainflow at station 1, and pc and V, are the density and velocity of t h e coolant at the coolant exits on the blades. P, and Me were used to calculate the denominator.

( 5 ) Tt T. =

1 + Y M : serves as an average isentropic exit temperature, and then

The numerator can be written as the massflow rate of the coolant, me, divided by the coolant exit area. Since the rotameter was calibrated for air in standard condi- tions, the volumetric flow rate read on the rotameter had to be modified for the coolant density at the meter.

Results Data were taken for each station of both cascades

in the Me range of 0.7 through 1.4 and with 0, 10,

20, and sometimes 40 psi blowing pressmes at the u p stream regulator. For a certain blowing pressure but different cascades and Mach numbers, the blowing rate was slightly different because the static pressure at the coolant exit was different. For this reason the blowing pressure is used for presenting some of the data and not the blowing rate; this latter was an output of the data reduction and could not be set prior to the tunnel runs. Approximate values of the blowing rates could still be attributed to the blowing pressures, and they are 0.4-0.6 for 10 psi, 0.8-1.2 for 20 psi, and 1.3-1.7 for 30 psi blowing pressures.

In Figure 5 the pitchwise distribution of isentropic Mach numbers at station 1 is presented for two different blowing pressures. The isentropic Mach numbers were obtained by using Eq.(l) with the local static pressure instead of the average exit pressure, P,. Little effect of the blowing can be observed.

M I I I I I . . . . I I I : . . . . i . . . , . , . . .

1.5

,.4 . . . . . .

..... i .... --S- B=O ..j ......... i ...._..... i ......... i ...._..... j ...... . . . . . . . . . . . , . .

Fig.6 Distribution of isentropic M a c h numbers at station 1. (B blades, M. = 1.14)

Some results of the blade surface pressure mea- surements for the B blades with no cooling applied are shown in Fig.6. The blade surface pressures were con- verted to isentropic Mach numbers in the same way as the wall static pressures were, and the distribution of these Mach numbers for different exit Mach numbers are plotted. The abscissa is the location of the mea- surement points given by z/c, that is their axial coor- dinate over the axial chord. For each exit Mach number two curves are given, one for the suction and one for the pressure side. The figure shows that as long as the flow was choked the suction side Mach numbers were independent of the exit Mach number up to the point where the oblique shock from the trailing edge of the blade above impinged on the instrumented blade. For the cases with injection the only thing that was possi- ble to measure was the effect of a cooled blade on the adjacent, not cooled blade. This again, was due to the fact that a blade could not be cooled and instrumented at the same time. Figure 7 is presented here for the B

1.6

1.4

1.2

1 .o

0.8

0.6

0.4

0.2

0.0

6.0

M

-~ - -

-

0 0.2 0.4 0.6 0.8 1 Fig.6 Blade surface isentropic Mach number

distribution (B blades, no cooling)

M I .6

1.4

1.2

1 .o

0.8

0.6

0.4

0.2

0 . 0 0.0 0.2 0.4 0.6 0.8 1 .o

Fig.7 Suction surface isentropic Mach number distribution (B blades, Me = 1.10)

blades with an exit Mach number of 1.23. The suction side isentropic blade surface Mach number distribution for 0 and 30 psi blowing pressure is shown. The effect of the blowing can be seen farther upstream of the throat, where the Mach numbers are lower for the cooled cases. This is due to the fact that the blowing decreased the effective throat cross section area.

stream and downstream of the blades Fig.6 iS presented for the time averaged exit Mach numbers of 1.34. The da ta shown belongs to the B blades, but the main fea- tures are similar for the U blades. The traversing took place at station 1, and no cooling was applied when the data for this figure were taken. The abscissa in this figure is the transverse, y coordinate, but it also could be time since the traversing makes a certain y location correspond to a certain time point. Therefore, even though Pru was measured with a stationary probe and we effectively measured it as a function of time, it could be converted into a function of y : 4. = P,,(y).

PI"- Ptl [psi1

,o,o , PlU. , Pll , [psi1 , , , , I ptu (lsiai 50

t

6.0

4.0

2.0

0.0

-2 .0

. . . . .

- B=0.85

1 Y / ' I 1

0 1 2 Fig.9 Traverses with different blowing rates

(B blades, station 1, Me,."- = 1.235) -'

To demonstrate the total pressure readings u p

6

A % 8 r I I I I

0 ' I I I I I I

0.7 0.8 0.9 1.0 1.1 1.2 1.3 1.4

Fig.10 To ta l pressure losses - (B blades, station 1) A %

A

0.7 0.8 0.9 1.0 1.1 1.2 1.3 1.4 Fig.12 Total pressure losses

(U blades, station 1)

A vo 8 I ! 1 ! I I

" 0.7 0.8 0.9 1.0 1 . 1 1.2 1.3 1.4 0.7 0.8 0.9 1.0 1.1 1.2 1.3 1.4

Fig.lS To ta l pressure losses (U b lades , station 2)

Fig.11 To ta l pressure losses (B blades, s t a t i o n 2)

is quite significant for this relatively high Mach number case. Also, it is the greatest between the wakes, where the Mach number is the highest. Furtheremore, this downstream traven reveals the good periodicity of the flow. The effect of the blowing can be seen in Fig.9,

The maSs averaged total pressure loss coefficients versus the time averaged exit Mach numbers are pre- sented in Figs.10-13. In each figure results with three or four different blowing pressures are included. The

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7

differences between the results for the B and U blades reflect the difference between their design Mach number and trailing edge thickness. The differencies between the results for station one and station two show that there is an additional losses due to mixing between the two stations. The differencies due to the different blow- ing rates while everything else was kept the same are not significant. There is a small trend noticable, that for low Mach numbers the cooling even reduced the losses, This effect can be explained by the momentum added with the cooling and that Eq.(3), which we used in the data reduction, did not account for that com- ponent. For higher Mach numbers, however, a slight increase in the losses is noticable. The experimental uncertainty was shown to be larger than these differ- encies, therefore one should be careful in interpreting these results.

Unce r t a in t i e s Extensive effort was made to determine the ac-

curacies of the results. The inaccuracies in the mea- sured variables were A p t , = 0.06 psi, A p e = 0.02 psi, A(Pt, -Pt2) = 0.02 psi, AQ, = 1 cfm and APd = 1 psi. From these the error in the calculated variables could be determined. The error in the exit Mach number was AM. = 0.005. Because of the error introduced by the upstream total pressure - wall static pressure cor- relation, the error in Me,,,, was sufficientally higher, about 0.02. The calculated error in the blowing rate was approximately 0.1.

The error in the mass averaged total pressure loss coefficient consisted of several parts. The errors in Pt, and in Pt, - Ptz caused 0.0015 error in A. The error due to taking the static pressure constant in the plane of traversing was also not greater than 0.0015. In spite of the relatively big changes in PI this error is quite low, and the reason for that was the cancelling effect of the integration. The third important contributor to the errors was the use of the approximate formula for cal- culating A , and this introduced an error no greater than 0.003. In fact, only in the case of the highest coolant - massflow ratios was the error this high. In most cases it remained below 0.002, and at higher Mach numbers its value dropped below 0.001. This third kind of error was the hardest to determine. Several assumptions had to be made, such as the assumption that the COz did not mix with the air, and it travelled downstream in a certain part of the wake. Finally, the correlation of the wall static pressures with the upstream total pres- sure introduced an erroT of about 0.003 in the worst cas-s With the neglection of other less important con- trlL, I os to the error of A , the combined error was estimated as Ad = 0.005 .

Conclusions

The presented results show that there was little effect of the coolant injection on the flowfield and on

the aerodynamic losses. Venediktov‘, who used similar geometry and blowing conditions, arrived at the same conclusion. Therefore, it can be concluded that with good design and placing of the trailing edge slots most of the losses generated by tangential coolant injection in the trailing edge region can be eliminated for reasonable blowing rates.

d

Acknowledgement This work was supported by General Electric Air-

craft Engines, Cincinnati , Ohio.

References 1. MacMartin, I. P., and Norbury, J. F.,”The Aero-

dynamics of a Turbine Cascade with Supersonic Discharge and Trailing Edge Blowing,” ASME pa- per No 74-GT-120, 1974.

2. Prust, H. W., ”Cold Air Study of the Effect on Turbine Stator Blade Aerodynamic Performance of Coolant Ejection &om Various Trailing Edge Slot Geometries, Part 11: Comparison of Experimental and Analytical Results,” NASA Paper No TM-X- 3190, 1975.

3. Lokai, V. I., and Kumirov, B. A., “Losses in Tur- bine Cascades with Cooling Air Discharge and Var- ious Trailing Edge Geometries,” Izw. Vuz. Aviats. Tekn., VOLE, No.3, 1973

4. Venediktov, V. D., ”Investigating a Turbine Stage With Cooling Air Leaving through Slots in the Concave Surfaces of the Nozzle Blades,” Teploen- ergetika, Vo1.19, No.7, 1972, pp. 15-19.

5. Lawaczek, O., ”The Influence of Jets of Cooling Air Exhausted from the Trailing Edges of a Su- percritical Turbine Cascade on the Aerodynamical Data,” AGARD CP 229, Paper No.30, 1977.

6. Zaccaria, M. A.,”Developement of a Transonic Turbine Cascade Facility,” M.S. Thesis, Dept. of Mechanical Eng., Virginia Polytechnik Institute and State University, 1988.

7. Singer, R. T., ”An Experimental Examination of the Effect of Trailing Edge Injection on the Aero- dynamic Performance of Gas Turbine Blades,” M.S. Thesis, Dept. of Mechanical Eng., Virginia Polytechnik Institute and State University, Decem- ber 1988.

8. Bertsch, R., “An Experimental Examination of the Influence of Trailing-Edge Coolant Ejection on Blade Losses in Transonic Turbine Cascades,’ M.S. Thesis, Dept. of Mechanical Eng., Virginia Polytechnik Institue and State University, Decem- ber 1990.

W