A I A A-9 4-2 5 0 9

Diffuser Modification Schemes Investigated for the NASA Ames 11-Ft Transonic Wind Tunnel

L. Goenka Sverdrup Technology, Inc. (until 7/92) Tullahoma, TN

This work resulted from studies performed under NASA contract NAS2-13032

18th AlAA Aerospace Ground Testing Conference

June 20-23, 1994 / Colorado Springs, CO - u

For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics 370 L'Enfant Promenade, S.W., Washington, D.C. 20024

Diffuser Modification Schemes Investigated for the NASA Ames 1 I-Ft Transonic Wind Tunnel

by

Lakhi N. Goenka Svcrdrup Technology, Inc.'

* until July 1992

Abstract

The NASA-Ames 11-Ft Transonic Wind Tunncl has a backleg diffuser system with significant now separation. This system consists of the post- comprcssor annular diffuser, and a wide-angle diffuscr located immediately downstream. The flow in the annular diffuser was separated for test-section Mach numbers below 0.8, while the flow in the wide-angle diffuser was always separated and lay io the jet-flow regime. Several improvement concepts were investigated in a 5.4 percent scale-model test facility a t NASA Ames called BLASTANE. The annular diffuser concepts included the use of vortex gencrators and radial flow-control rods, while those for thc wide-angle diffuser included annular vanes, radial splitter-plates, large-scale vortex generators (LSVG's), and a mid-diffuser screen. The radial now-control rods (mounted on the nacelle in the annular diffuser), in conjunction with the annular m m , resulted in an improvement in both pressure rccovery, as well as flow stability, of the backleg diffuser system, although one quadrant in the wide- :ingle diffuser still exhibited separated flow. This problem was later resolved in subsequent tests (not rcported hcre) by the addition of a screen downstream of the annular vanes.

L

The radial flow-control rods provided a uniqne solution to improving the flow in the annular diffuser, and sewed as a guide for improved annular ditliuscr fixes used in subsequent NASA tests. They produced an exit velocity profile which was rclatively insensitive to the variations in the diffuser inlct profile. Further, these rods did not generate a significant total-pressure loss. Therefore, they helped improve the flow stability, without compromising

the total-pressure recovery, of the backleg diffuser system.

1.0 INTRODUCTION

The NASA-Ames 1 I-Ft Transonic Wind Tunnel is used extensively for scale-model tests in the transonic and supersonic regimes. NASA recently embarked on a project to improve the flow- quality in this tunnel. This included the minimization or elimination of low-frequency oscillations observed in the test section of this tunnel. Earlier tests on the tunnel had indicated that these low-frequency oscillations resulted from large- scale flow separations in the tunnel's backleg diffuser. Such low-frequency oscillations could adversely &wt test results, and an effon was undertaken to define diffuser modification schemes to minimize or eliminate the flow separation in the backleg diffuser.

The backleg diffuser system consists of an annular diffuser located downstream of the compressor, and which contains the compressor nacelle, as well as a wide-angle conical diffuser that immediately follows the annular diffuser. Earlier tests on the 11-Ft Tunnel had shown that the flow in the annular diffuser separates under certain tunnel operating conditions, while the flow in the wide- angle diffuser is always separated.

The diflber improvement concepts aim at eliminating the flow separation in these diffusers and improving the flow distribution into the aftercooler, which is located at the wide-angle diffuser exit. It

was hopcd that besides minimizing the low- frequcncy effects observed in the test section, some improvement in the total-pressure recovery of these diffusers would result from these modifications, realizing a savings in tunnel power consumption. Also, some improvement in test section temperature uniformity may be obtained because of improved flow into the aftercooler.

The results of scale-model tests on two annular diffuser and four wide-angle diffuser modification concepts are presented in this paper (Ref. I ) . These tests were conducted in a 5.4-percent scale-model facility, which is described in subsequent sections. Some of these concepts were based on cstablished wide-angle diffuser [modification tcchniqucs, while others represented innovative solutions to solving the diffuser problem. I t should bc noted that the concepts presented in this papcr arc not the final diffusermodification concepts selectcd, but hclped provide a basis for defining the latter. Dctails of the final concepts selccted will appear i n other NASA dcvelopcd publications.

I \

2.0 ANNULAWIDE-ANGLE DIFFUSER GEOMETRIES

A schematic of the 1 I-Ft Transonic Wind Tunncl is included in Fig. 1. The 1 I-Ft Tunnel is a closcd-loop, transonic wind tunnel with a maximum test-scction Mach number of 1.5 and a maximum

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total prcssure of 32.5 psia. As shown in Fig. I, the annular and wide-angle diffusers to be modified are located in the backleg of the tunnel, between the compressor and the aftercooler.

The geometry of the annular diffuser is depicted in Fig. 2. It consists in part of a cylindrical duct which is 69.33-A-long and 2 4 4 in dinmcter. A

17.00 A protrudes into the cylindrical duct to from the annular diikser. Five, 1 5-in.-thick tailconc

53.63-A-long tailcone with an inlet diameter of W

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Fiw-1. cro-arrdt*arrr,s-.

support struts are located immediately downstrcam of the last row of compressor exit vanes. These supports are 11.33-A long. The tailcone section in this region is conical, and its diameter reduces to 16.67 A at the exit of the support struts.

W The location of the annular diffuser on a

diffuser map is shown in Fig. 3. This dilfuser has an area ratio of 2.0, and an equivalent conical half- angle of 3.725 deg (assuming that its length is equal to that of the tailcone, which is 54.33 R). This puts it close to the diffuser stall-line, and makes the flow within the diffuser sensitive to non-uniformities in the inlet velocity profile. The map in Fig. 3 also includes annular diffuser sections of other wind tunnels for comparison.

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2

The wide-angle diffuser is located immediately downstream of the annular diffuser. A schematic of this diffuser is presented in Fig. 2. It has a wall angle of 30 deg, an area ratio of 8.5, an exit diameter of 70 A, and a length of 39.5 A. The aftercooler is located at the wide-angle diffuser exit, and has a flow-deflection screen mounted on the ccntral portion of its upstream face to deflect the flow out towards the aftcrcooler panels nearthe walls.

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The esreme geometry of the wide-angle diffuser causes large-scale flow separation that bcgins at the diffuser inlet (jet-flow regime). Thc resistance of the aftercooler is not suRcient to keep thc flow within the diffuser attached. This flow scparation increascs the diffiser total-pressure loss, and caiiscs the mean-flow unsteadiness, as well as possibly the tempcraturc non-uniformities, in the test sccLion.

NASA recently ran tests in the 1 I-Ft Tunnel to characterize the flow in the 1 I-Ft annulat and wide-angle diffusers. The following general conclusions can be drawn from these tests:

0) The flow in the annular diffuser varies with tunnel operating conditions. There is a velocity defect, and possibly some scparation, at the outer walls of the annular diffuser at tcst-section Mach numbers below about 0.8. The situation reverses itself at higher Mach numbers, when there is separation off the tailcone, with a resultant velocity defect at the tunnel centerline.

Increasing the Reynolds number has a beneficial effect on the flow in the annular diffuser. There is a reduced tendency for flow separation at a given Mach number as the tunnel total-pressure is increased.

Thc flow in the wide-angle diffuser is completely separated at all test section Mach numbers, and lies in the jet-flow regime. There is little or no diffusion (ie., there is no static-pressure recovery) in the flow within the wide-angle diffuser.

The combined loss across the diffuser system, including the flowdeflection screen and the cooler, is about 0.5 qe, where Q is the compressor-exit dynamic pressure.

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( i i )

(iii)

(iv)

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3.0 NASA-AMES BLASTANE SCALE- MODEL TEST FACILITY

Tests lo validate the diffuser modification concepts were conducted in a 5.4 percent scale- model facility at NASA Ames, called BLASTANE. A schematic of this facility is shown in Fig. 1. This

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facility draws in air from the atmosphere, and is driven by a vacuum pump attached to its downstream end. The annular and wide-angle diffuser scctions, as well as the 'aftercooler section and the flow- deflection screen, were modeled during these tests. Details such as the existing nacelle contour, as well as the nacelle support struts, were also included. The aftercooler was carefully simulated using perforated plates, and its resistance coefficient and geometry were also replicated.

The walls of the annular and wide-angle diffuser were made from lexan for good flow visualization;' - f i is enabledvery good visual access of the d i f l k r flowfield. Tufts and tnft wands were used for flow visualization. 0.25-in. diameter swivel fittings were installed in the diffuser walls to permit the insertion of the tuft wands.

The measurements made to quantify the diffuser pressure recovery included wall static- pressure measurements made along the bottom wall of the diffuser system. Total-pressure rake surveys at five different cross-sections of the diffuser flodield were used to obtain velocity distributions.

The two typical compressorexit profiles which are obtained in the 1 I-Ft Tunnel (as described in the preceding &ionsj were simulated in the NASA-Ames BLASTANE facility. The profile for test-section Mach numbers below 0.8 (which exhibits a velocity defect along the outer wall of the annular

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diffuser) was simulated using an inlet perforated fence and air bleed through a small circumferential

. . ~ ~ . . +1 - L 1 gap in the ducting. No such treatment was required 1 6 j .

to simulate the profile for test-section Mach numbers above 0.8. It should be noted that none of these techniques was successful in reproducing the flow- separation phenomenon observed in the full-scale

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I 0.84 tunnel. The diffiser velocity profiles, as well as wall-static pressure data, obtained for the "baseline" 0 5 j ..

0 41

0.2, '

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configuration -- i.e. with no diffuser fixes, are presented in Figs. 5 and 6. Because the compressor- exit profile for test-section Mach numbers below 0.8

flow-separation in the diffuser system, the diffuser improvement concepts described is the following scctions were at first tested for this case. Test results for only this inlet profile are reponed. Successful

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represents a much more severe case for inducing i .0 0.1 0 2 0:3 014 0.5 0.8 0.7 08 09 1 0 rlR

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k i ~ r 0 . 8 i ~ t h ~ l 1 - f t T - d k difluscr fives obtained were, however, tested with the

adcqnate for this condition. othcr compressor-exit profile to verify that they were 12.0-

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4

1.n ANNULAR-DIFFUSER MODIFICATION CONCEPTS

The annular diffuser modification concepts were considered to eliminate the separation in this diffuser, and to improve the flow at the diffuser exit.

Two different annulardiffuser modification

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concepts were investigated. These concepts were considered to eliminate separation in this d f i s e r , and to improve the flow at the diffiser exit. These concepts include: (0

(ii)

Vortex Generators on the diffuser outer walls Radial Flow-Control Rods on the tailcone of the annular diffuser

1.1 Vortex Generators

Vortex gcncrators have been used succcssfully in conical diffusers having half-angles of up to 8 degs (Senoo and Nishi, Ref. 2). They cncrgizc the boundary-layers, thereby suppressing flow separation along the walls of the diffuser. It was hoped that this action would improve the flow in the annular diffuser. There were, however, two complicating factors in the present application: ( I ) The diffuser inlet profile varied with tunneI operating conditions. The defect in the inlet velocity profile near the diffuser walls implied that this problem may not be confined to just the flow boundary-layers. (2) The wide-angle diffuser section immediatcly follows the annular diffuser, so that the flow had to undergo further diffusion. This implied that the fix should generate a flow having enough momentum near the outer walls when entering the wide-angle diffuser, and that it would not be adcquate to just suppress the separation in the annular diffuser.

b

The vortex generators tested consisted of flat plates, and had a height of 12 in., a chord of 12 in . , and angle of attack of 12 deg. Tests were conducted using forty such vortex generators, spaced iiniformly on the diffuser outer wall, and located about 2-ft downstream of the trailing edges of the first row of tailcone support struts. These vortex generators were tested in both counter- and co- rotating configurations.

4.1.1 Test Results

The results of the BLASTANF. tests indicated that +e vortex generators were unsuccessful in improving the flow in the annular diffuser. .They had negligible effect both on the pressure recovery of the annular diffuser, as well as on its exit velocity profile. This is because the defect in the velocity profile at the compressor exit (Le. entering the annular diffuser) was not confined just to the boundary layer, but extended ont well into the outer flow - i.e., it did not represent a boundary- layer problem.

4.2 Radial Flow-Control Rods

Radial flow-control rods (or stars) have been used successfully to improve the pressure recovery and flow stability of moderately wide-angle two-dimensional and conical diffusers. The use of these rods to improve the flow in twodimensional diffusers was at first investigated by Kmonicek (Ref. 3). Eight rods were mounted radially on a hub to form a star-shaped device, which was mounted on a central rod and inserted into the twodimensional diffuser. Flow stability and pressure recovery were increased with particular rod combinations, while a sudden discontinuity in flow stability was observed when the rods were slightly withdrawn from a side wall. This concept was then extended to conical wide-angle diffusers. The stars showed significant promise, increasing both pressure recovery and flow stability for diffuser wall angles ranging from 7.5 to 15 deg. Further investigations on the use of radial flow-control rods to improve the performance of conical diffusers with an half angle of 1 1 deg and an area ratio of 4 were conducted by Welsh (Ref. I). The investigations by Welsh revealed that thc maximum diffuser pressure recovery was dependent upon the rod thickness, star location, and diffuser inlet conditions. For diffusers with a half angle of 11 deg and with inlet velocity profiles having blockage fractions of about 0.02, a star should be located at d r i= 1.9 and have an area blockage of 10 to 14 percent, where x is the axial distance from the diffuser inlet, and rj is the diffuser inlet radius. For inlet velocity profiles with blockage fractions of about 0.1, a star should be located at x/ri= 1.9 and have an area blockage of 20 percent. The overall diameter of the star should be equal to the diffuser inlet diameter.

5

A significant distortion in the diffuser exit flow arose from the use of the radial flow-control rods. The exit velocity profile was double-peaked, with a large overshoot at the outer walls. The d i f i se r exit flow was also highly turbulent. It was rclt that these two qualities would be beneficial to the performance of the wide-angle diffuser in the present application. Also, the double-peaked velocity profile is inherently unstable, and should therefore mix rapidly a short distance downstream.

The radial flow-control rod geometry (full- scale dimensions) used is shown in Fig. 7 . Eight rods, each having a diameter of 16.5 in., were located at station 1123.5 in., which is 49 .634 downstream of the compressor exit. Note that the compressor exit is at station 528 in. These rods had an c\-lerior diameter of 17 R, which is the cquivalent diameter (dc,) of a conical diffuser having the same arca as the inlet to the annular diffuser. The local area-blockage of these rods was 14.7 percent, with tlic tailcone diameter at this station equal to 5.54 A.

The radial flow-control rods were located much further downstream (drei= 4.9) than their suggcsted optimum location in a conical diffuser ( d q = 1.9). There was, however, no information available as lo the optimum location of these rods in :in annular diffuser. It was felt that locating the rods fiirthcr upstream wonld ( I ) make them too short and stubby and (2) result in their being located at a station at which the diameter of the tailcone is disproportionately large; this would decrease their cffcctiveness.

/

4.2.2 Test Results

LJ The BLASTANE tests demonstrated that the use of radial flow-control rods on the tailcone was a partly successful annular diffuser fix. The effect of the% rods on the velocity distribution in the diffuser system are shown in Fig. 8 (a) & (b). They were successful in redistributing the flow momentum towards the outer walls for the case when the annulardiffuser inlet velocity profile exhibitcd a defect at its outer walls. This was because their local frontal-area blockage increases towards their rmt (i.e., towards the tailcone). When the inlet velocity profile exhibited an overshoot at the diffuser outer walls, the flow-control rods did not appreciably affect the flow in the annular diffuser. As a result, they produced an exit velocity profile which was insensitive to the variations in the diffuser inlct profile.

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Diffuser C!&i 3). Ob- ria Rm-CwQd Rods L. AamuI.~ Dimur.

The effect of the flow-control rods on the pressure recovery of the diffuser system is shown in u

6

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improvement was obtained in subsequent tests by the addition o f a middiffuser screen at the exit of the annular vanes. The results of these tests are not presented in this paper, but will be described in other NASA publications. The pressure recovery of the diffuser system with just the annular vanes was dCp= +0.28. Therefore, the pressure loss associated with the flow-control rods when used with the annular vanes was only dCp= -0.02. However, large-scale flow separation in two quadrants of the wide-angle diffuser was observed.

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5.2 Other Wide-Angle Diffuser Concepts

The other wide-angle diffuser concepts investigated included radial splitter-plates. large- scale vortex generators, and a middiffuser screen. These concepts were not successful in improving the flow stability within the wide-angle diffuser, and will therefore be only briefly described below.

5.2.1 Radial Splitter Plates

v Radial splitter-plates divide the diffuser into

smaller compartments with reduced equivalent conical angles. This concept is described in Ref. 6, which utilized eight splitters in a 38deg total angle diffuser with an area ratio of 15. A significant improvement in diffuser effectiveness was obtained, with a disc placed at the apex of the radial splitter- plates (and serving as a separation trip) proving to be critical to the success of this concept. Further investigations on this concept can be found in Refs. 7 and 8. The splitter-plate concept used for the present wide-angle diffuser modifications is shown in Fig. i i . T i e plates have a iength of i6 ieet, and a radiai disc having an area blockage of 2 percent is located at their apex. Full-length splitter-plates were not used since the resistance of the aftercooler was e x w e d to fi l l the wide-angle diffuser once the flow

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exited the splitter-plates.

The BLASTANE tests showed that the radial splitter-plate concept proved to be ineffective. The plates were not successful in spreading the flow to the dif€user outer-walls, despite the fact that central discs with blockages up to 10 percent wcre

8

used. Wide-angle difiser velocity profiles were not significantly altered by the splitter-plates

1 I

Dim"- b Wbt#

11. Cramdry 01 Radial Splinrr-PlSIa

(see Fig. 12). This was perhaps because the low- cnergy flow in the nacelle wake impaired the flow- sprcading action of the central disc.

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5.2.2 Large Scale Vortex Generators

Large Scale Vortex Generators (LSVG's) are pyramid-shaped inserts that are attached to the expansion-wall of the wide-angle diffuser. LSVG's were developed by the author (Refs. 9 & IO) in a wide-angle, plane-wall diffuser with a wall-angle of 20 deg and an area-ratio of 2. These devices differ fmm conventional vortex generators in that they modify the inviscid flow in the diffuser, while conventional vortex generators only modify the boundary-layer flow. The LSVG geometries tested are shown in Fig. 13 (a) and (b). The first configuration consisted of six LSVG's spaced equally along the circumference of the diffuser i n k . The upper surface of these inserts was swept up so thc apex of the pyramid was raised up one-third the distance to the diffuser centerline. The projectcd frontal-area of the upswept insert was designed to be the same as the low-upsweep insert tested in Refs. 9 and 10. The second configuration consisted of four inserts spaced equally along the circumference of the wide-angle diffuser inlet. The length of these inserts wasabout half the length of the wide-angle diffuser, and was therefore closer to the full-length configurations tested in Refs. 9 & IO. The apex of each pyramid was raised up one-half the distance to the diffuser centerline.

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9

separation in the wide-angle diffuser, and had little or no effect on the diffuser velocity profiles.

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~i~~ ixb). ~ t o a r m , o r ~ . m S a k V o * C = = c n m n (C.m.np"h 1).

The BLASTANE tests indicated that the LSVG's were not successful in controlling the flow- scparation in the wide-angle diffuser. Though they had some effect on the dif€user velocity profiles, the cxit flow was grossly asymmetric (Fig. 14), implying that all the inserts were not successful in generating longitudinal vortices. This was probably the result of thc extremely rapid rate of change of the area along the length of the wide-angle diffuser. A dip in the wall static-pressure at the wide-angle diffuser inlct indicated the presence of a local low-pressure region induced by the formation of a vortex on one of thc LSVG's (Fig. 14 (b)).

5.2.3 Mid-diffuser Screen

Thc use of screens to fill wide-angle diffusers is wcll known, and is documented in Refs. I I & 12. Tlic resistance and location of the scrccn(s) can be obtained using the formula in Rcf. I I . The screen used had a loss coefficient of 1.8, and was located 6 feet downstream of the inlet to the wide-angle difFuser. This screen configuration assumcd that the loss-coefficient of the aftercooler (which scrvcs as a second screen) was 25, as indicated by earlier NASA tunnel characterization tcsts.

The BLASTANE tests showed that this scrccn was iinsucccssful in climinaling the

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6.0 SUMMARY AND CONCLUSIONS

The NASA-Ames 1 I-Ft Transonic Wind Tunnel has a backleg diffuser system with significant flow separation. This system consists of the post- compressor annular diffuser, and a wide-angle diffuser located immediately downstream. Earlier NASA tunnekharactenzation tests had shown that the compressor exit-velocity profile vaned with tunnel operating conditions. The flow in the annular diffuser was separated for t e s t - d o n Mach numbers below 0.8, while the flow in the wide-angle diffuser was always separated and lay in the jet-flow regime.

Several improvement concepts for the backleg diffuser system were investigated i n a 5 I

v

10

percent scale-model test facility at NASA Ames called BLASTANE. The annular diffuser concepts includcd the use ofvortex generators and radial flow-control rods, while those for the wide-angle diffuscr included annular vanes, radial splitter- platcs, large-scalc vortex gcnerators (LSVG's), and a inid-diffiiscr scrccn.

- Tlic scale-model tests indicated that the

radial flow-control rods (mounted on the nacelle in the annular diffuser), in conjunction with the annular vanes, resulted in an improvement in both prcssurc rccovery, as well as flow stability, of the bncklcg diffuser system, although one quadrant in the wide-angle diffuser still exhibited separated flow. This problem was later resolved in subsequent tests (not rcportcd lierc) by the addition of a serecn downstream of thc annular vanes.

The radial flow-control rods provided a uiiiquc solution to improving the flow in the annular difkscr, and scrvcd as a guide for improved annular diffuscr fizcs used in subsequent NASA tests (not rcportcd hcre). They were sricccssful i n rcdistribriting the flow momcntum towards the diffiiscr ontcr walls for the case when the annular- dif iser inlet velocity profile exhibited a defect at its outcr walls. This was because their local frontal- area blockage increases towards their root (i.e., towards thc nacelle). When the inlet velocity profile csllibikd an overshoot at the diffuser outer walls, the flow-control rods did not appreciably affect the flow in thc nnnnlar ditruser. As a result, they produced iiii csit vclocity profilc which was relatively iiiscnsitivc to the variations in the diffuser inlet profilc. Further, these rods did not generate a significant total-pressure loss. Thereforc, they hclpcd improve the flow stability, without compromising the total-pressure recovery, of the backlcg diffuser system.

v

Acknowledgments

I would like to thank Mr. Lado Muhlstein and Mr. Kcn Mort o f the NASA Ames Research Center for their rcview o f the initial test plan and reports, as wcll as for their support during the NASA Ames BLASTANE tests. I would also like to acknowledge thc significant contributions made by Alan R. Boone, Fraiik J. Kmak, Sreedhara V. Murthy, and Dong E. Muzrio, both in the planning as well as the implementation of the NASA Ames BLASTANE lcsts.

W

Table 1. Dimensions of Annular Vanes (see Fig. 9).

Vane No. Inlet Exit Vane Radius Radius Angle

(A) (fo (deg)

1 6.00 10.40 16.09 2 8.50 11.72 22.18 3 10.4 18.00 26.48 4 12.00 20.80 30.00

References

1.

2.

3.

4.

5.

6.

Boone, Alan R., Kmak, Frank J., Murthy, Sreedhara V., and Muzzio, Doug E., Private Communication with respect to NASA Ames BLASTANE Test Results, NASA Ames Research Center, Moffett Field, CA, 1992.

Senoo, Y. and Nishi, M. "Improvement of the Performance of Conical Diffusers by Vortex Generators." ASME Paper No. 73- WA-FE-I, June 1973.

Kmonicek, V. "Flow Through a Conical Diffuser with Flow-Stabilizing Star Shaped Insert." Zprava 42/56, 1956.

Welsh, A. M. "Flow-Control in Wide- Angled Conical Diffusers." Journal of Fluids Engineering, December 1976.

Turner, J. T. '"Improvement of Wide-angle Conical Diffuser Performance by Means o f Conical Vanes." Simon Engineering Laboratories, The University of Manchester.

Rao, D. M., and Raju, K. N. "The Use of Splitters for Flow Control in Wide-angle Diffuses." National Aeronautical Laboratory, Bangalore, India. TN-AE-26- 61, 1964.

11

7 . Rao, D. M. "Potential Application of Radial Splitter Ditfuser to Shrouded Wind Turbines." Journal oJEnergy, Vol. 1, No. 2, February 1977.

Rao, D. M. "A Method of Flow Stabilization with High Pressure Recovery in Short, Conical Diffusers." The Aeronautical Journal, Vol. 75, May 1971. pp. 336-339.

Gocnka, L. N., Panton, R. L., and Bogard, D. G. "Studies of Flow Patterns in a Diffiser Designed to Generate Longitudinal Vortices." Journal oJFluids Engineering, Vol. I 11, June 1989.

Goenka, L. N., Panton, R. L., and Bogard, D. G. "Pressure and Three-Component Velocity Measurements on a Diffuser Designed to Generate Longitudinal Vortices.'' Journal of Fluids Engineering, September 1990.

Schubauer, G. B. and Spangenburg, W. G. "Effect of Screens in Wide-angle Diffusers." NACA Report 949, 1948.

Mehta, R. D. "The Aerodynamic Design of Blower Tunnels with Wide-Angle DilTiisers." Prog. Aerospace Sci., 1977, Vol. 18, pp 59-120.

8.

9.

IO.

1 I .

12.

12