NASA / TMm2001-210674

Effect of Delta Tabs on Free Jets

from Complex Nozzles

K.B.M.Q. Zaman

Glenn Research Center, Cleveland, Ohio

National Aeronautics and

Space Administration

Glenn Research Center

March 2001
Acknowledgments

The experiments with the ACE nozzle model were planned and conducted in consultation with Dr. J.M. Seiner of

NASA LaRC who is closely familiar with the full scale.

Note that at the time of research, the NASA Lewis Research Center

was undergoing a name change to the

NASA John H. Glenn Research Center at Lewis Field.

Both names may appear in this report.

Available from

NASA Center for Aerospace Information

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Available electronically at http://gltrs._c.nasa.gov/GLTRS
EFFECT OF DELTA TABS ON

FREE JETS FROM COMPLEX NOZZLES

K.B.M.Q. Zaman

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Introduction:

The effect of tabs, or small protrusions in the flow placed at the exit of a nozzle, on the

evolution and spreading of free jets has been under investigation at NASA Lewis Research

Center. The tabs introduce streamwise vortices in the flow which can incwa_ the jet spreading

significantly. The streamwise vorticity generation mechanism, and the resultant effect on

spreading, noise, thrust loss, etc., for axisymmetric jets have been studied in detail and the

results summarized in Ref. 1 (see also, Refs. 2-4).

With non-axisymmetfic, e.g., rectangular or elliptic, nozzles the effect of the tabs

becomes relatively more complex. 5'6 It is found that, depending on the nozzle geometry and

placement of the tabs, jet spreading can either be increased or decreased. In such jets streamwise

vortices are usually already present due to secondary flow within the nozzle. The vortices

produced by the tabs then interact with the already present vortices and, depending on the sense

and the strength of the resultant vortices, the subsequent evolution of the jet may be affected

differently. The phenomenon of axis switching in such jets can also be stopped or promoted

depending on the placement of the tabs. _ These observations underscore the powerful role of the

streamwise vortices in the evolution of free jets.

On the other hand, it should be borne in mind that practically all the vorticity shed

initially from the nozzle is azimuthal. Even with the tabs, the intensity of the azimuthal vorticity

at any downstream location usua_y far exceeds the intensity of the streamwise component; the

peak of we is typically measured to be about 5 times larger than the peak of _x .5 Thus, any

process which organizes the azimuthal vorticity and leads to the formation of coherent vortical

structures can also have a profound effect on the jet evolution. The influence of the dynamics of

the streamwise and azimuthal vortices and their interaction on jet spreading is currently being

further investigated. 7

While carrying out the aforementioned studies of a fundamental nature, efforts have also

been made simultaneously to apply the tabs to more complex nozzles of practical interest. In

NASA/TM--2001-210674 1
support of the High Speed Research (HSR) program, experiments have been carried out using

simplified models of nozzles that are candidates for the High Speed Civil Transport (HSCT)

aircraft. The purpose of the present report is to summarize and document the results obtained

with the latter nozzles.

A primary goat in the HSCT program is jet noise reduction with acceptable performance

loss. All nozzle configurations considered in the program utilize ejectors. A basic concept in

such a configuration is to mix the primary jet with the lower speed outer stream so that a lower

velocity at the end of the ejector is obtained. For a given thrust, the lower exhaust velocity

results in lower jet noise. The penalty, of course, is larger and heavier nozzle hardware.

Therefore, a goal in the design of such a nozzle is to mix the primary jet as fast as possible so

that the length and the weight of the ejector can be reduced. Improved mixing within the ejector

could also result in additional noise benefit through removal of 'hot spots', i.e., isolated pockets

of high velocity fluid. It was thought that the tabs might be helpful in achieving such goals, and

this led to the present experiments.

In terms of jet noise reduction, the tabs could have further beneficial effects. They could

reduce screech associated noise -an effect that has been well known. 8 Also, the tabs could be

applied at the end of the ejector to increase the spreading of the resultant jet exhaust. It is

possible, even though not completely clear, that this might produce some direct noise benefit.

That an increased spreading of a free jet produces less noise has been indicated by the analytical

study of Ref. 9. The experimental results on the effect of tabs on far field noise of axisymmetric

jets, reported in Ref. 10, also lend support to this notion.

The aim in the present experiments has been to look for and demonstrate increased

mixing with models of the HSCT primary nozzles through the use of tabs. The bulk of the

experiments have been of a trial and error nature, in view of the flow complexities even with the

simplified models of the nozzles. The results do show some promise. Only the more promising

tab configurations are investigated in detail in regards to jet spreading, noise and thrust loss, and

are reported in the following. Obviously, the effectiveness of the tabs cannot be confidently

predicted, and tests must be conducted with full scale hardware which will be even more

complex involving hot flow and the ejector. The main objective in the present effort has been to

come up with recommendations for tab configurations for possible full scale verification.

Models of two nozzles have been considered. One is the lobed mixer nozzle used in most

of the configurations considered for the HSCT. The other is an axisymmetric plug case

simulating the 'ACE' (Axisymmetric Coannular Ejector) nozzle which is also a candidate for the

HSCT. Because of the basic differences in the two nozzle shapes, the results are presented in

two parts. Part A documents data obtained with several models of the lobed nozzle. Part B

documents those obtained with the 'ACE' nozzle model with various designs of the plug.

NASA/TM--2001-210674 2
Nomenclature:

Cfl -- Measured thrust normalized by ideal thrust obtained from measured mass flux and NPR,

by assuming convergent, uniform flow.

Cf2 - Measured thrust normalized by ideal thrust obtained from measured mass flux and NPR,

by assuming ideal expansion.

D - Equivalent diameter based on nozzle exit area.

d -- Diameter of end hole on a porous plug.

- Mass flux at a given x.

• _ -- Mass flux at nozzle exit estimated for a given NPR assuming uniform flow.

Mj - Jet Mach number had the flow expanded fully for a given NPR.

NPR -- Nozzle pressure ratio (Pt/pa).

Pt - Plenum chamber pressure.

Pa - Ambient pressure.

T = Jet thrust; 'ideal' is isentropic prediction assuming convergent, uniform flow.

U,V,W - streamwise, transverse and spanwise mean velocities.

u',v',w' - turbulent normal stresses.

uv, uw - Turbulent shear stresses.

Uj = Jet velocity at nozzle exit.

x,y,z -- streamwise, transverse and spanwise coordinates.

_x = Mean streamwise vorticity, _V/0z- 8W/_.

Lobed nozzle notations: six-lobed #1 has an end wall at the exit, six-lobed//2 has thin lip and

outside chutes cut out.

Plug notations: 'xx%H' or 'xx%S' means plug with xx% surface porosity with 'H' for circular

holes providing the porosity and 'S' for slots providing the porosity. A total of five plugs with

10 ° half-angle are used: one is solid, one is with slots and 10% porosity, others are with holes

with porosities of 3.6%, 5.5% and 8.2%. 'Ace_8.2%H_05' denotes nozzle with 8.2% porous

plug and d = 0.05 inch end hole.

Tab notations: 'nn.a%' denotes equally spaced nn delta tabs with a total area blockage of a%.

Three tab configurations are used with 6, 12 and 24 tabs, each with a total area blockage of 4%.

Two other configurations are used with 6 and 24 tabs, each with a total area blockage of 1%.

'12._4%' tab denotes 12 delta tabs with 4% total area blockage.

NASA/TM--2001-210674 3
Experimental Facifities:

The experiments were conducted in the jet facilities housed in ERB CW13 and CW17.

All nozzles studied were convergent ones, converging from a round inlet to the specific exit

geometries; a length of approximately 1.27 cm with constant cross section was allowed before

the exit. All jets discharged into the ambient of the test cell. Laboratory air supply was used and

the flows were continuous with feedback control for holding a set Mach number. Flow field

measurements were taken by automated probe traversing mechanisms under computer control.

The smaller jet facility in CW13 was used for much of the earlier work on the tab effect

which have been reported in Refs. 1 and 2. In short, this facility has a 11 cm diameter plenum

chamber supplied with compressed air with maximum available tank pressure of 80 psig (560

kPa). Nozzles used with this facility have equivalent diameters, D - 1.47 cm. Mass flow rate

could be measured with an orifice meter fitted to the supply line. Jet thrust could also be

measured using a load cell and an auxiliary plenum mounted on linear bearings. All flow

visualizations and jet noise measurements were done in this facility.

The larger jet facility is housed in cell CW17. It has a 76 cm diameter plenum chamber

fitted with flow conditioning units and an acoustic treatment section. A maximum plenum

pressure of 125 psig (875 kPa) and flow rates more than 3.6 kg/sec are available. All nozzles

used in this facility for the present experiments have equivalent diameter, D = 6.35 cm. The

hot-wire measurements for the flow fields were done at a nominal Mach number, _ = 0.31.

Limited data were also obtained in this facility at supersonic conditions.

The hot-wire measurements involved use of two crossed wire probes, one in the u-v and

the other in the u-w configurations. The probe traversing was done such that the two probes

went through the same grid points on a cross sectional plane at a given downstream location.

The measurements provided distributions of U, V, W, u', v', w', fi'_ and fi'_. Gradients of V

and W yielded streamwise vorticity, e x. Standard Pitot probe measurements were carried out.

The Pitot probe data for the supersonic jets, obtained far enough downstream where the flow

was subsonic, provided the Mach number distributions on cross sectional planes. Integration of

these data provided streamwise mass fluxes. Further details of the measurement techniques can

be found in Refs. 1 and 3.

PART A LOB/_ M/XER NOZZLE

Fig. AI: The exit section of the large jet facility is shown in (a); the coordinate system used in

this report is also indicated. The shape of a 'delta tab', fitted to a circular nozzle, is shown in

(b). This is a shape which was reached after some parametric study to produce maximum flow

field distortion for a given area blockage. Unless otherwise stated, a 'tab' will imply a 'delta

NASA/TM--2001-210674 4
tab' with dimensions w/D = 0.28 and _ = 135 °. The corresponding area blockage per tab is

approximately 1.7 % of the nozzle exit area.

Two geometries for the lobed nozzle were tested. The smaller, six lobed nozzle had D =

1.47 em and its exit geometry is shown in (c). Two versions of this nozzle were used. One (#1)

had a cylindrical outer shape with a diameter of about 5.7 cm; the face of the cylinder at the

exit plane formed an 'end wall'. A retainer disc was attached to the end wall to hold the tabs.

The second one (#2) had exactly the same interior dimensions as #1, however, the outside was

shaved so that the nozzle ended with a thin lip. Thus, this nozzle had the outside chutes cut out

through which there was some secondary flow via natural entrainment. The six lobed nozzles

were fabricated by EDM.

The other, larger, eight lobed nozzle had similar exit dimensions as a Pratt & Whitney

nozzle (NASA Contract No. NAS3-25952; see also Ref. 11). The model had D = 6.35 era, and

only the exit dimensions were to scale to approximately simulate the primary flow of the P&W

nozzle but the interior shape was arbitrary. The interior converged from 13.3 em dia round

shape to the exit shape; the exit dimensions are shown in (d). This nozzle also had an end wall

for easy installation and location of the tabs. In the following, this nozzle is referred to as the

P&W nozzle. It was fabricated by stereolithography.

Fig. A2: Two views of the P&W nozzle are shown. In the lower picture the retainer can be seen

holding one tab on each side.

Fig. A3: Mach number contours for a supersonic flow at x/D = 14, measured with a Pitot

probe in the large jet facility, are shown in this figure. Data for the P&W nozzle are compared

with those from a circular nozzle fitted with four delta tabs. The circular nozzle had the same

exit area (D = 6.35 cm).

The contrast is dramatic. An underlying concept for a lobed nozzle is to stretch the

mixing layer region exposed to the ambient so that there is more entrainment and spreading.

However, it is clear that in regards to that objective the action of the tabs with the circular jet

outperforms the lobed nozzle under consideration. In fact, the normalized mass flux for the

lobed nozzle was found to be smaller than that for the circular nozzle even without the tabs. The

normalized flux values are indicated by the numbers in parentheses on the lower center of these

figures. The flux data are further discussed with the next figure. It is noteworthy here that four

delta tabs (of the given size and dimension) with the circular nozzle were found to be the

optimum for increasing jet spreading.l Equally spaced delta tabs with numbers either more or

less than four resulted in less spreading.

NASA/TM--2001-210674 5
Fig. A4: The variation of the normalized mass flux with strearawise distance was measured in

the small jet facility. Data for the circular jet with and without four delta tabs are compared with

those for the six lobed nozzle case. The two solid data points correspond to the large jet case

shown in fig. A3 ('cross' for the P&W nozzle, 'plus' for the circular jet with four delta tabs).

Even though the normalized fluxes were somewhat lower for the larger nozzles, it is clear that

the data trends are essentially independent of Reynolds number which varied by almost a factor

of 5 between the two facilities (Re D = 4.2x10 6 for the large jet). Also, as mentioned in the

foregoing, the fluxes for the lobed nozzles are found to be consistently lower than even the

baseline circular jet case.

The result that the lobed nozzles produce less jet spreading compared to the circular

nozzle came as a total surprise. Eventually the reason became apparent. At the given Mach

number the circular jet went through a screech, whereas there was no screech with the lobed

nozzles. (Noise spectra for the six lobed nozzle will be shown in comparison with the ACE

nozzle data in Part B). A screeching jet is a periodicatly forced jet in which mixing enhancement

can take place by an organization of the azimuthal vortical structures. Such an effect has been

demonstrated by acoustic forcing of subsonic jets in several previous studies (e.g. Ref. 12).

Thus, the presence of screech with the circular nozzle and the absence thereof with the lobed

nozzles is believed to have caused the anomaly noted in the foregoing.

As expected, data for the same nozzles of fig. A4, at a subsonic condition (not shown

here for brevity), do show a reversal of the trend with the fluxes for the lobed nozzle exceeding

that for the circular case. The fluxes for the circular jet with four delta tabs, however, remain

by far the largest even at the subsonic condition. Note that the nozzle configuration with the

delta tabs in figs. A3 and A4 does not involve any screech and the mixing increase occurs

through a different mechanism, i.e., through the dynamics of the streamwise vortices. It may be

of interest to note here that at Mj = 1.63 the circular jet goes through a screech in the helical

mode (mode C). 13Screech in the flapping modes (B and D), at higher or lower values of Mj,

results in a further increase in jet spreading. These results will be further analyzed and discussed

in a later publication.

It should be noted here that the circular jet with four delta tabs is one of the few

configurations, found so far, which result in such a large increase in jet spreading. The fluxes

obtained are larger than those obtained with the lobed nozzles or non-axisymmetrie (e.g.,

elliptic) nozzles. This is true for supersonic as well as subsonic conditions. Also, at subsonic

conditions the increase in the fluxes caused by four tabs far exceeds that achieved by acoustic

excitation. Comparison with data for elliptic jets (Refs. 14,15) and acoustic excitation (Ref. 12)

has been discussed in Ref. 1.

NASA/TM--2001-210674 6
Fig. A$: The lower flux values measured with the lobed nozzle at first led to the notion that the

blockage due to boundary layer thickness in the narrow channels could be high. If that were the

case the initial mass flux would be lower, causing the anomaly, as t_, the normalizing quantity,

was calculated assuming uniform flow (zero boundary layer thickness). However, if this were

the case jet thrust would also be correspondingly lower. The thrust data shown in fig. A.5

disproved this notion. Both versions of the six lobed nozzle produced thrusts only slightly lower

than that expected from the 'ideal' which was also calculated assuming uniform flow.

The thrust data obtained by using six tabs with the lobed nozzle is also shown in this

figure. The tab locations are indicated and the sizes are given in the following with the

discussion of fig. AT. At NPR = 4.47 0VIj = 1.63), the thrust loss from 'ideal' is found to be

about 1.6% for the no tab cases and 6.5 % for the six tab case. It should be noted that these

thrust loss numbers are overestimates of the performance loss. With the tabs there is flow

blockage and for the same NPR the actual mass flow through the nozzle is lower. The appropriate quantifier in this connection is the coefficient of performance (cfl or cf'2). These

measurements were done for the ACE nozzle and will be diccussed in Part B. The coefficients

cfl and cf2 were not measured in the experiments with the six lobed nozzles. However, a rough

estimate for the percent performance loss could be obtained by subtracting the percent area

blockage from the percent thrust loss. Not counting the blockage due to the boundary layer

thickness, this would result in a thrust loss of about 3.5% for the tab case.

Fig. A6: The tabs were tried with the six lobed nozzle in an effort to increase mixing. That the

tabs significantly alter the flow fields could be demonstrated readily by laser sheet flow

visualization, as shown in this figure. The visualization technique is based on natural moisture

condensation in the mixing layer where ambient air comes in contact with the cold core of the

supersonic jet. Thus, the mixing layer region is illuminated by the laser sheet and seen in these

pictures. It should be clear that four tabs, placed as shown in the sketch at the top of the right

column, distort the jet differently compared to the no tab case.

Fig. AT: Similar flow visualization was conducted for various tab configurations, and the results

are shown schematically in this figure. Tabs of two sizes were used: one had w - 0.28 cm

representing approximately 0.8% area blockage, the other had w ffi 0.18 cm representing

approximately 0.34% area blockage (w is defined in fig. A1). The mixing layer outline as seen

at x/D = 3 are shown. Note that the tabs placed in the inner channel, as represented by the third

case from the top, make little difference on the flow field. Also, a remarkable difference occurs

depending on the size of the tabs. The tabs, on the outer four channels, are large in the fifth

case from the top but are small in the sixth case from the top. This difference caused the

different lateral spreading of the jet as shown.

NASA/TM_2001-210674 7
Fig. A$: The mass flux variations obtained with the sixth case of fig. A7 are compared with the

flux data shown before (fig. A4). For the sake of completeness, data obtained with both versions

of the six lobed nozzle are shown in this figure. Recall that version #2 had outer chutes cut out

so that there was some secondary flow due to natural entrainment. Clearly, this did not increase

the mixing; in fact, the fluxes were even lower. The reason for this difference remained unclear.

The tabs, however, certainly made an improvement. The thrust loss for the case with the tabs

was discussed with fig. A5.

The rest of the data in Part A are for the P& W nozz/e. Only flow field measurements

could be conducted with this nozzle due to facility constraints. The large nozzle size, however,

allowed adequate spatial resolution for the hot wire measurements. 5'7 Key results, obtained by

hot-wire as well as Pitot probe surveys, are presented in the following.

Fig. A9: The data shown in fig. A3 were for _ = 1.63 which corresponded to a flow rate of

about 3.6 kg/sec. Such a large flow rate caused high recirculation velocities in the limited

volume of the test cell (approximately 12m x 6m x 6m). The flow noise was also too high to

permit experimentation during normal working hours. Thus, a lower Maeh number was chosen

for further studies at supersonic conditions with the P&W nozzle.

Data, similar to those shown in fig. A3, for different tab configurations are shown in this

figure at Mj = 1.2. The tab configurations are indicated by the insets in each figure. All tabs

had the same size with w ffi 1.1 cm (fig. A1), which corresponded to an area blockage of about

0.7% per tab. Clearly, the effect of the tabs here is not as dramatic as observed with the circular

jet. Nevertheless, some increase in the fluxes and changes in the jet cross sectional shape are

caused by the tabs.

Fig. AIO: Perhaps, of more significance is the effect of the tabs nearer to the nozzle exit which

would be pertinent to the flow within the length of the ejector (typically a few equivalent

diameters from the primary nozzle exi0. Maeh number distributions at x/D = 2, for the eases of

fig. A9, are shown in this figure. The effect of the tabs in the near flow field is quite pronounced. The two and six tab cases make the jet cross section fuller which would be desirable

with the ejector. Referring back to fig. A9, note that the eight tab configuration produced the

largest increase in the mass flux downstream. However, this configuration may not be as

effective with the ejector as the jet spreads mostly laterally at the upstream locations.

Fig. All: Data corresponding to the cases of fig. A10, for _ ffi 0.31, are shown in this figure.

'Whereas all the flow field data shown in the foregoing were obtained by Pitot probe traverses,

the data in this and the rest of the figures in Part A were obtained by hot wire anemometry.

NASA/TM--2001-210674 8
When compared to the data of fig. A10, the effect of the tabs can be found to be quite

similar at supersonic and subsonic conditions. A similar observation was made earlier with

axisymmetric nozzles, which led to the inference that compressibility has little to do with the

primary effect of the tabs._ This is reassuring as the hot-wire measurements, conducted to

determine the details of the flow field, could be carried out only at low subsonic conditions. It is

likely that the inferences drawn from these data, e.g., about the vorticity dynamics, should be

generally applicable at all Mach numbers.

Pig. A12: In this figure the streamwise evolution of the jet cross section for the no tab case is

shown. The peaks and valleys in the velocity distribution are seen to smear out by x/D - 2.

Fig, A13: The variation of the mean velocity along the axis of the nozzle is shown in (a) for the

four tab configurations. Corresponding variation measured with a circular nozzle is also shown

for comparison. A faster decay of the jet velocity is indicated for the lobed nozzle, and this is

affected differently by the different tab cases. A complete picture regarding the jet evolution can

only be obtained by the detailed velocity distributions over the entire volume of the flow field.

However, the centerline variations are easy to obtain and can be used for cursory assessment of

the jet evolution. Such presentation is also customary in the literature and this prompted the

inclusion of fig. A13. The corresponding centerline variations of turbulence intensities are

shown in (b). Of note is the fact that the case with six tabs results in lower intensifies and also a

significant upstream shift of the second peak. The data shown in (a) also indicate that this case

involves the fastest overall decay in the mean velocity.

Pig. A14: The evolution of streamwise vorticity, corresponding to the data of fig. A12, are

shown in this figure. Each lobe produces a pair of counter-rotating streamwise vortices. Thus,

initially there are eight pairs of such vortices. These, however, go through amalgamation and by

x/D = 2 there are only four pairs left. A similar observation was also made earlier with circular

jets; when too many tabs were used the vortices would undergo interactions and settle down to a

fewer number. It is as if the jet cross section can accomodate only about two pairs of streamwise

vortices and this number would be reached no matter how large the number is initially. The two

resultant pairs of vortices, however, have been found to persist for a long distance downstream. 5

Fig. A15: The effects of the different tab configurations on the streamwise vorticity distribution

are shown in this figure. Clearly, there occurs a profound redistribution of the streamwise

vorticity under the influence of the tabs.

NASA/TM---2001-210674 9
Fig. A16: The sense of rotation of the streamwise vortex pairs observed with the P&W nozzle is

explained in this figure. Each lobe produces a counter-rotating pair with the 'out flow' sense of

rotation. 7 These vortices are produced due to secondary flow within the nozzle. 16 Placement of a

tab produces a vortex pair with opposite sense, as shown in the inset on the left. 1 The vortices

from the tab can be so strong that they simply overwhelm the naturally occurring vortices.

However, the strength of the vortex pair depends on the tab size, a larger tab produces a

stronger pair. Thus, since vorticity is diffusive, it might be expected that there exists a certain

smaller size of the tab which will just null the naturally occurring vortices and there would be

no resultant pair downstream.

When the tab is placed on the side of the nozzle, as shown in the inset on the lower

right, it produces a vortex pair that augments the naturally occurring vortices occurring from the

nearest lobes. For example the vortex on the right corner from the upper lobe has the same sign

as the upper vortex produced by the tab. These two would amalgamate and produce a reinforced

vortex. Such reinforcement of the natural vortices explains why just two tabs placed on the sides

produced a remarkable effect on the flow field.

Fig. A17: Some of the detailed flow field data obtained together with the data of figs. A12 and

A14 are documented in this figure. The logitudinal normal stress is shown in (a) while the two

components of shear stresses are shown in (b) and (c). The contour interval is 0.02 in (a) and

0.001 in (b) and (c).

Fig. A18: Corresponding turbulent stress distributions for the tab cases of figs. All and A15

are shown in this figure. Not much further could be inferred on the dynamics of the flow field

from the data of figs. A17 and A18. These data are included as they might be helpful in

possible future analysis and computation of these flows.

Concluding remarks on Part A:

In terms of increasing jet mixing the effect of the tabs appears to be more pronounced

than that of the lobes. The primary effect of the tabs is to produce streamwise vortex pairs while

that of the lobes is to stretch the area of the mixing layer exposed to the ambient. The present

data indicate that mixing increase due to the latter mechanism is not as efficient as that due to

the dynamics of the vortex pairs. However, many lobe geometries are designed to produce

strong streamwise vortices which might be just as effective as the tabs. Furthermore, the lobes

of the present nozzle produced streamwise vortices through secondary flow, and thus, the

relative roles of 'area increase' and 'streamwise vortex dynamics' in mixing enhancement remain

far from completely clear. This issue as well as the dynamics of the s_wise vortex pairs

responsible for increased mixing need to be addressed in the future.

NASA/TM--2001-210674 10
The investigation also indicates that the jet flow field can accomodate only a few pairs of

streamwise vortices. Thus, if too many vortex pairs are generated initially the vortices go

through amalgamation or diffusion and settle to a fewer number. This occurs within a few

equivalent diameters from the nozzle exit. Based on this evidence the usefulness of too many

tabs or too many lobes might be questionable.

With the lobed nozzles, application of the tabs certainly alter the flow field. The effect is

especially noticeable within the first few diameters from the nozzle exit. Based on the data

obtained with the P&W nozzle it is thought that the cases involving two and six tabs have the

best chance in producing positive effects when applied to the full scale experiment. Such an

experiment with an actual Pratt & Whitney nozzle, together with the ejector and hot flow, is

currently underway in the Nozzle Acoustic Test Rig (NATR).

PART B 'ACE NOZZLE" MODEL

Background: These experiments were conducted in the small jet facility with a simplified model

of the ACE nozzle which is another configuration being considered for the HSCT aircraft

(NASA Contract No. NAS3-25963). The nozzle model used in the present experiments approximately replicated only the primary flow. In a first set of experiments, a solid, pointed plug

with 13 ° half-angle was used. Subsequently a 10° half -angle plug with rounded end and

variable surface porosity was used. The effect of the tabs together with the different plugs was

studied. Flow visualization, noise and thrust measurements were conducted. Because the tabs

had to be applied to a narrow annulus in this ease they had to be of smaller size compared to the

size indicated in fig. Al(b). Two sets of delta tabs were used. In each set the area blockage to

the flow was held constant while the number of tabs was varied. An objective of the investigation was to assess the effect of the number of tabs on thrust loss, mixing, and noise from the

jet. The motivation came from an earlier observation with round nozzles which showed that, for

a given area blockage, fewer larger tabs performed better than many smaller tabs. Key results

are presented in the following.

Fig. BI: Pictures of the flow facility, nozzle and tabs are shown. The top picture shows the

auxiliary plenum which is supplied with compressed air by four radial flexible tubes. The tubing

seen in the middle is for measuring the plenum pressure. The plenum is mounted on linear

bearings to enable thrust measurement with the help of a load cell.

The picture in the middle shows the nozzle fitted with the pointed plug. The picture at

the bottom shows the five tab discs. The tab discs were installed with a retainer and screws. The

NASA/TM--2001-210674 11
tab discs were wire-cut from shim stock with thicknesses ranging from 0.0076 cm to 0.025 cm

depending on the sizes of the tabs. The tabs were bent approximately 45 ° with a suitable die.

The tab designations are indicated in the figure and explained in the nomenclature. The base

width of the tabs for the five cases, reading from left to right in the figure, is 0.264, 0.188,

0.137, 0.137, and 0.066 cm. The last one, the 24_1% case, is approximately comparable in

scale to the hardware used in an earlier Boeing experiment.

Fig. !12: The nozzle fitted with the pointed plug is shown schematically in (a). The nozzle

converged from 3.81 cm dia to 1.91 em dia then diverged and converged over the centerbody

(plug) terminating with a diameter of 2.1 cm. The shape and location of the plug was such that

the flow converged throughout the annular passage. Thus, all the supersonic cases involved

underexpanded flows. The annulus at the exit was measured with filler gages and found to be a

maximum of 0.292 cm on one side and a minimum of 0.279 cm on the other; the equivalent

diameter, D was 1.44 era. The outer lip of the nozzle was not sharp and the thickness (i.e., the

'end wall') was left to allow easy installation of the tab discs.

The geometry of the 10 ° half-angle, rounded plug is shown in Co). The porous part of

the plug is interchangeable. The axial location of the plug was adjusted so that the equivalent

diameter of the annular exit was about the same as with the pointed plug. The porous plug

notations have been explained in the nomenclature section.

Fig. 113: This figure shows the laser sheet illuminated cross section of the jet about 0.6 cm

downstream from the tip of the pointed plug. As stated earlier, the visualization shows the

mixing layer region as the bright area. The asymmetry in the jet cross section is believed to be a

result of the camera angle and the three struts holding the plug (fig. B2a). It can be seen that

some of the tab configurations clearly increased the jet spreading. More spreading occurred with

the cases involving lesser number of tabs. The 24 tab case in both sizes (1% and 4% blockages)

produced minimal effect.

The rest of the results presented in the following are obtained with the 10 ° rounded

plugs.

Fig. !14: Sample flow visualization pictures with the rounded plugs are shown in this figure.

(These pictures were obtained at a different time and, in contrast to fig. 133, the flow is from

left to right in these pictures.) Essentially the same effect of the tabs was observed. The effect of

the porosity of the plug was not discernible from these pictures.

Fig. 115: Schlieren photographs are shown for the flow field around the plug. The shocldexpansion structure around the plug can be seen for the no tab cases. The structures appear to be

somewhat diffused with the 5.5%H plug compared to the solid plug ease. Again, the difference

NASA/TM--2001-210674 12
in the upper and lower halves in these pictures is believed m be due to the asymmetry introduced by the struts holding the plug. The effect of the tabs in either plug case is profound, as

seen in the lower two pictures. The discrete shock/expansion structure appears to be weakened

and diffused significantly.

The effect of 6 and 24 tabs, with the solid plug, is shown in fig. BSCo). The effect of 6

tabs is similar to that of 12 tabs. However, here some discrete shock structures with reduced

spacings can be discerned. With 24 tabs the shock structure approximately returns to the

structure seen for the no tab case.

Thrust Loss:

Fig. 116: This figure shows comparison of thrust (T, Newtons) vs. NI'R for the four porous plug

cases with that for the solid plug ease. All eases are seen to involve significantly lower thrust

from the 'ideal'. This is likely to be due to substantial boundary layer thickness relative to the

annulus dimension. 'Ideal' thrust implies isentropie calculation assuming uniform flow at the

nozzle exit (zero boundary layer thickness). The thrust variations are similar for all plug cases.

At NPR = 4.47, the thrust loss from 'ideal' is found to be 5.7% for the 3.6%H case and 6.9%

for the 10%S case, the others being close to 6%.

Note that this and the rest of the figures in the report are presented in their original

computer printout forms. The tables on the right may be ignored. However, a curious reader

might find some more information from these. For example, the percent thrust loss at NPR =

4.47 is listed in the upper table. In the lower table, '%blkg' represents (two times the) percent

flow blockage computed from measurements of actual mass flow, the nozzle pressure ratio and

the nozzle exit area; 'ed0' represents orifice meter discharge coefficient obtained from calibration with a circular nozzle; 'cfl' and 'cf2' are coefficients of performance described in the

nomenclature.

Fig. B7: The effect of the end hole size (see fig. B2b) on thrust is shown. The hole diameters in

inches are indicated. The effect, for the present case, is negligible.

Fig. 118: Thrust versus NPR is shown for the indicated tab cases. All three tab configurations

involve thrust loss, the least is with the 6_4% tab case.

Fig. 119: Thrust coefficient vs. NPR are shown corresponding to the data of fig. BS. Here, the

six tab case exhibits the least performance loss. With the pointed plug, twelve tabs performed

best (data are not shown for brevity). However, uncertainty in these data are large, estimated to

be equal to one ordinate division, due mainly to error in the mass flow measurement.

NASA/TM--2001-210674 13
Fig. B10: While the data in fig. B9 were taken with the solid plug, this figure shows corresponding data taken with the 5.5%H plug. Here, the loss due to the three tab configurations is

pmcticaUy identical.

Noise data:

Far field noise was measured at a distance of 67D, and an angular location of 110 ° from

the downstream axis of the jet. Spectral analysis was done with an analyzer (Nicolet 660B) with

bandwidth _f = 250 Hz. It should be noted that the test cell was only semi-anechoic and thus

the amplitudes may be contaminated especially at the lower frequency end. However, the

comparisons for the relative effects of the tabs or the porous plugs should be valid.

Fig. Bll: Spectra for the four porous plug cases are compared with that for the solid plug case.

Successive pairs of curves are staggered by one major ordinate division; the ordinate shown

pertains to the pair at the bottom. In (a), at the higher NPR, the porous plugs are seen to make

little difference in the noise spectra compared to the spectrum for the solid plug case. This is not

so at the lower NPR in (b), where all porous plugs are seen to produce additional noise!

Again, the notations and tables on the right of these figures may be ignored. But

additional information can be obtained; the list in the table pertains to the spectra traces starting

from the bottom. For example, in CoL rec 307 is the solid curve and rec 306 is the clashed curve

in the bottom pair. The fully expanded Mach number (Mj) is indicated in the second column;

'scr_f' represents frequency (kHz) of the highest peak in the spectra, 'scr_a' represents the

corresponding amplitude in dB, and 'oaspl' represents overall sound pressure level in dB.

Pig. B12: Data for three tab configurations together with the 5.5%H plug are compared with

data for the no-tab, solid plug case. The 12_4% tab results in the most overall noise reduction,

at both NPR.

Fig. B13: Data for the porous plug and the solid plug are compared. Together with the tabs, the

porous plug results in slightly lower amplitudes at the higher NPR in (a). The comparison is

much more favorable for the porous plug at the lower NPR in (b).

Fig. B14: Data for the porous plug and the solid plug, without tabs, are compared in (a) at four

values of NPR. For some unknown reason (and there are many in this complex problem!), the

solid plug produces less noise at NPR ffi 3.5 (see also fig. B1 lb). At other values of NPR, the

difference in the data is little.

The data in Co) show that the slot plug (dashed curves) produces additional noise

compared to the plug with holes (solid curves), especially at lower NPR.

NASA/TM--2001-210674 14
Fig. B15: The effect of 12_4% tab case (with the solid plug) is compared with the no tab case

at four values of NPR. At higher NPR, the reduction in noise at lower frequencies is accompanied by some increase at higher frequencies.

Fig. B16: The effects of the 12_4% tab with the solid plug (dashed curves) and the 5.5%H plug

(solid curves) are compared at four values of NPR. The porous plug yields some additional

noise benefit.

Fig. 17: Finally, the noise spectra for the ACE nozzle, with and without the tabs, are shown in

comparison to the spectra for the six lobed nozzle (Part A) and for an elliptic nozzle. The two

pairs of spectra at the top show comparison between the six lobed nozzle and the 3:1 elliptic

nozzle, on both minor and major axis planes. As mentioned with fig. A4, no screech is observed

for the six lobed nozzle. The ACE nozzle without tabs can be seen to have screech comtmnents.

The pair of spectra at the bottom shows that the ACE nozzle with the 12_4 % tabs produces the

least noise among the cases considered in this figure.

Concluding remarks on Part B:

The following inferences are made:

(1) Jet spreading is undoubtedly more with fewer larger tabs.

(2) Tab effect on noise far outweighs the effect produced by the porous plugs.

(3) Together with the tabs, the porous plugs produce additional noise benefit over the solid plug.

(4) The 12-4% tab together with the 5.5% porous plug produced the least noise.

The reason why many small tabs, as compared to few larger tabs, do not work well even

with a simple round nozzle is not completely clear. It is possible that, since a tab does not work

well if it is submerged in the boundary layer, an increasing ineffectiveness with smaller size may

explain this trend. For a given total area blockage, or for a given size of each tab, there should

exist an optimum number which would work best. For about 1.7 % area blockage each, four

delta tabs with a simple round nozzle increased the jet spreading the most; increasing the

number of tabs caused a reversal in the effect. For

expected to be larger.

In the ACE nozzle model it is apparent that

smaller size tabs the optimum number can be

the boundary layer thicknesses are large

relative to the width of the annulus. This might be the reason why the smallest size tabs did not

make a noticeable effect on the flow field. In the full scale nozzle the boundary layer can be

expected to be much thinner relative to the annulus width. This may be why a scaled up tab

configuration corresponding to the 24_1% ease was quite effective in an earlier Boeing

experiment. However, based on the present results, six to twelve delta tabs may be expeeted to

work even better. This ought to be tried in the full scale experiment. It is also desirable that for

a given number of the tabs, the total area blockage be varied, perhaps, ranging from 0.5% to

6%, to find the optimum effect.

NASA/TM--2001-210674 15
References:

tZaman K.B.M.Q., Reeder M.F., and Samimy M., 1994, "Control of an axisymmetric jet using vortex

generators', Physics of Fluids A, 6 (2), pp. 778-793.

2Ahuja K. K. and Brown W. H., 1989, "Shear flow control by mechanical tabs', A/AA Paper No. 89-0994.

3Samimy M., Zaman K.B.M.Q. and Reeder M.F., "Effect of tabs at the nozzle lip on the flow and noise field of

an axisymmetric jet', A/AA J., 31(4), pp. 609-619 (1993).

_Reeder, M.F. and Zanum, K.B.M.Q., "The imp_t of 'tab' location relative to the nozzle exit on jet distortion',

30th Joint Prop Conf., Indianapolis, IN, July 27-29, 1994, A/A/I 94-3385.

SZaman, K.B.M.Q., 1993, "Streamwise vorticity generation end mixing enhancement in free jets by 'delta-tabs",

AIAA Shear Flow Conf--, July 6-9, 1993, A/AA 93-3253 (NASA TM 106235).

eT.mnan K.B.M.Q., 1994, "Effect of 'delta tabs' on mixing and axis switching in jets from asymmetric nozzles"

AIAA Paper 94-0186, Aerospace Sciences Meeting, Reno, NV, Jan.

_Zaman K.B.M.Q., 1995, "Axis switching and spreading of an axisymmetric jet - role of vorticity dynamics',

A/AA Paper 95-0889, Aerospace Sciences Meeting, Reno, NV, Jan., 1995.

S'l'anna H.K., "An Experimental Study of Jet Noise, Part II: Shock Associated Noise', J. Sound & V/b. vol. 50,

429-444 (1977).

_Dahl M.D. and Morris P.J., "Supersonic jet noise reductions predicted with increased jet spreading rate',

submitted for ASME/JSME Fluids Engineering Conf., Hilton Head, SC, Au 8. 13-18, 1995.

t0 Kobayuhi, H., Oinuma, H., Sawamura, T. and Outa, E., "Effects of tab size on supersonic underexpanded cold

and heated jet noise suppression and jet thrust loss', AIAA Aeroacouatics Conf.. (93-4348), Long Beach, CA, Oct.

1993.

tt Wolter J.D. and Jones C.W., "Pratt & Whitney two dimensional HSR nozzle test in the NASA Lewis 9-by 15-

foot low _eed wind tunnel: aerodyna_c results', Proc., First NASA/Imtustry HSR Program NozT.le b_#nposium,

Nov 17-19, 1992.

nCrow S.C. and Champagne F.H., "Orderly structures in jet turbulence', J. Fluid Mech. , 48, 547-591 (1971).

tSNomm T.D., "Screech suppressi(m in supersonic jets', A/AA ./., 21, pp. 235-240 (1983).

t4Ho, C.-M. and Gutmark, E., 1987, "Vortex induction and mass entrainment in a small_-ratio elliptic jet _, J.

Fluid Mech., vol. 179, pp. 383-405.

tSHussain F. and Husain H.S., 1989, "Elliptic jets. Part 1. Char_teristics of unexcited and excited jets', J. Fluid

Mech., 208, pp. 257-320.

teBradshaw, P., "Turbulent secondary flows', Ann. Rev. Fluid Mech., 19, pp. 53-74, (1987).

NASA/TM--2001-210674 16
EXPERIMENTAL FACILITY

(a) Large jet and coordinates

r

= 41

(b) delta tab

Diameter = D

---I.M

(c) Six lobed nozzle (d) Eight lobed (P&W) nozzle

6.1

l

FIG. A1

]7
EIGHT LOBED 'P & W' NOZZLE MODEL

IM

¢-M-MTI+

18
M-CONTOURS; x/D = 14, Mj = 1.63

4.

o

o.

l I I I I I l

- . O. 2 4

I I I I I I I

0.92

0.04

(2.38)

___. ' O I. ' 21. '

z/D

FIG. A3

NASA/TM--2001-210674 19
NORMALIZED MASS FLUX VERSUS x/D, M i = 1.63

1

at

m

u

10.

I I I I I I I

C),

©

I I I I

_. _. _. _ .

x/i)

FIG. A4

NASA/TM--2001-210674 20
THRUST VERSUS NPR

100. I I I I I I I

T (N)

'Ideal'

#1

/

/

O!

I

!

3 I. 4 I.

Pt/Pa

FIG. AS

NASA/TM--2001-210674 21
LASER SHEET FLOW VISUALIZATION, r¢_ = 1.63

FIG. A6

NASA/TM--2001-210674 22
SKETCHES OF JET DISTORTION AT x/D., 3

FOR INDICATED TAB CONFIGURATIONS, Mj = 1.63

FIG. A7

NASA/TM--2001-210674 23
NORMALIZED MASS FLUX VERSUS x/D, Mj - 1.63

g

6.

r /r e

_['l m

m

1(

I I I . ! I I I

m

#1

FIG. A8

NASA/TM--2001-210674 24
M-CONTOURS; x/D = 14, Mj --- 1.2

4.

B

OQ

-2.

O.

__._o -

! I i I [ I I

0.56

O4

(3.06)

' I ' I ' I '

I I I I I I I

0.46

04

(3.42)

-)2. ' o_. ' 2'.

I

i

I

I

I

I I I I I i I

0.5 0.O4

(3.13)

' l ' I ' I '

| I | I | I I,

0.40

O.O4

(3.91)

!

- _. ' 0 I. ' 2 I. 4.

z/D

FIG. A9

NASAfI'M--2001-210674 25
M-CONTOURS; x/D = 2, Mj = 1.2

].. _ I I I I I I I I ! I

0.96 - 0.96

0.5-

-0. 5-

.08

O.

_0°

-I. -i. 5 ' 01 5 ' 015 ' i. -0 I.5 0.15 i. 5

z/D

FIG. A10

NASA/TM--2001-210674 26
U/UcCONTOURS;x/V = 2, Mj = 0.31

1. ¢_"

0. 5 m

--0. 5_

Oo 5_

--0. 5-

--1°5

5

0.1

I I , I. I

I _ I

L l I

\

O.9

-o! s ' o.1_ 1 5

I

'1

I

°

|

.--I

I

I

i

--4

i

I

-I

, I

, h , h , I , I I

..,, °1

0.8

I I I j I

0.9

' I I O" I

--0. 5 5

0.1

i

I. 3

zlD

FIG, All

NASA/TM--2001-210674 27
U/Uj-CONTOURS; FOR INDICATED x/D, Mj - 0.31

e

1. S

O. 5--

-0. 5-

Oo S_

-0. 5--

0.9

t z . t i "I I l i I I i

/

x/D=0.5

' I ' t' '

I I I I I

0.65

0.1

q

I

x/D=2

' I ' I '

x/D=l

-Z. 5 ' I ' 0) 5

-1.5 -0.5

0.9

I I l I I

0.8

X

IQ -o I.5 i.

z/D

FIG. AI2

NASA/TM--2001-210674 28
CENTERLINE MEAN VELOCITY VARIATION, Mj = 0.31

•

O.

u/uj O.

O.

Ol

1,

x/D

FIG. A13(a)

NASAFI'M--2001-210674 29
CENTERLINE TURBULENCE INTENSITY VARIATION

ffi 0.31

u_,uj

OI

I I I I I

, .,! , _! , _.!.

x/v

FIG. A13(b)

NASA/TM--2001-210674 30
to.D/U£CONTOURS; FOR INDICATED x/D, Mj - 0.31

1.5

0.5--

--0. 5-

--1.5

t I l I l

1.2

_..:'0 _'-'' i),.

IIIW, I-IT __

x/D=0.5

I ' '1

I I I 1

-1.2

1.04

4;:.

V

-,,...,:,_,..

.5

x/D=1

I !

-o 1. 5 o. [ 5

I

4

I

I

q

I

-1

..-I

I

1. 5

I I I I I

O.4

....::)_:),

x/D=2

' I ' I '

I I I I I

O.24

.....::::::]

_ffD=3

' / ' I '

--0. 5 O. 5 1 5

z_

FIG. A14

NASA/TM--2001-210674 31
¢_,D/U£CONTOURS; x/D = 2, Mj = 0.31

O° 5--

, -O. 5-

0.5--

-O. 5-

--1. _"

--1 5

I I I I I

,p,,,_,?.4,.i ,:,.,_ _.:@_

":"':'.'S"." \

-0.4

I I

_..2_,,., P..'&N,

L, ,_&,',

I I I

:.e_&,.

' oL ' - 5 O. I 5 5

I I I I t

_;'_,:,:." .'_k::""

':::::i ......'i..-"

0.44

; I ; I ;

I I _ I I

4' ,..:.__: ._#..__t.',_

"" __ •.,. o.-. -%

t"-, _ ',.. ",I ...."_.-?....

;:.-:::,'_ '_."_ ....

x

] o.,

' 01 5 ' 01_ ' I

z/D

5

FIG. A15

NASA/TM--2001-210674 32
SKETCH OF STREAMWISE VORTICITY DISTRIBUTIONS

/

FIG. A16

NASA/TM--2001-210674 33
u'/Uj-CONTOURS; FOR INDICATED x/D, Mj - 0.31

O° 5--

-O. 5-

-1.5

1.5

--O. 5--

5

t ! I i ] ....] , i t I I

x/D=0.5

[ I [ t

x/D=2

I I 1 i

I I I I I I I I I

0.13

/

1 oo . 11

x/D=l

-0! =_ ' O. I 5 1, 5

0.13

x

x/D=3

--0. 5 O. 5 i

z/D

FIG. A17(a)

NASA/TM--2001-210674 34
d'_/Uj2-CONTOURS; FOR INDICATED x/D, Mj - 0.31

e

1.5-

--O. 5-

--0. 5-

--1-5 I

I J I t t | I I I I _

J I

x/D=0.5

I |"' I -=

O.005

O.006

,,-::-. .. ,'.--.

",C_L-%':_--!:>i ) ..... Z:,'.':',," j

x/D=2

i' ]" I

J I

I I I I I I I I I ...... I

O.OO6

l I

i

I

- --!

. --4

l

l

5

O.005

i

:._==.-.:::: ........... ::::::::::::::::::::::::::::

_ _ ",....--'" ., .,.

;.... ,:...... .. c ..... ,.<',.'"")",'i ("

• _,..... ,° ? - _, ._-. ,.

x/D=3

-ol._ ' ;.'_ '

z/D .

FIG. A17(b)

NASAITM--2001-210674 35
ff_/U/-CONTOURS; FOR INDICATED x/D, Mj -- 0.31

1.5

--0. 5--

1. =

0.5-

--0. 5-

x/D=0.5

' I ' I '

I I I I 1

,',_

?II:

if::/I,

W

-O.OO5

!.

x/V-1

I I

-o I.'_ o.I

II --

x/D=2

I ' I

I ol

I

i

I

I

I

.--I

I

I

,,-i

I

I

5

I I I I

..,".,,'..-"?\...' ii"..,.,"//

.... ....,....,__.... ,/( ._

! /,/./ / ; _, % /

""-..'"-L ..)1///'_\ _ d ".'

.... ,,,.,-- /i"_(_ ' _]

',_;:-v ,.,.-

x/D=3

--OL 5 ', O. I 5

zfl)

FIG. A17(c)

NASA/TM--2001-210674 36
u,_cCONTOVRS;x/D - 2, _ - 0.3_

1.5

--O. S_

0.5--

--0. 5--

5

I I I I I

0.13

' I ' I '

i i I

/

0.13

I I

, -- j ° 5 '

I

0.5

I I I l i

0.13

I ' I ' I '

I I I I ! ,,

5 ' 005 0 5 1

5

z/D

F_G.MS(a)

NASA/TM--2001-210674 37
fi_/Uj:..CONTOURS; x/D - 2, Mj - 0.31

1,5

O_ 5--

--O. 5-

--1. _-

3..5

O° 5--

--O. 5-

; I ' I

I I I I I

5

I I

-01 5 O. 1.5

-7

----4

I

J

'I

I

"I

I

i

i 5

z/D

I I I I I

I ' I '

I I I I I

0.OO7

•================================================.

I I I

-0 I. _ O. I 1 5

_c,.. A18(b)

NASA/TM--2001-210674 38
_-_/Uj2.CONTOURS; x/D = 2, Mj -- 0.31

e

1.5

O° 5_

-O. 5-

-l.=_ i

1. =

0.5---

--O. %-

5

I I I I

•' o.' , I

.,;,:,:,.'.. ,_.o,v, i ::-'

,',.,;_;'.;,?.,;,',.--,.

",',', ':. '," "''.' Z',,i

I 1 I I I I

O.0O6

' I '

I I I

..,..-., / °.o. _

"."". i /," ; _

..,:.o i;::

II ' " . "/i ,..,,,'-.-.'..'

,2: ,;,_" ,....

!.ll, I I;I

';_i_ !!.'}

,.....: .. _ " ,"::'!i;i_,

£ :" ": r/ f'w_,'

",;.L.._7., i',"

I I

I I

I

-o I. 5 o. I

.,'..---, ! ,.";,".....; ';

//,.;.//[ _ _i(...../,':.)..::'Y

: : \.,.,,.," _-......

\ ._.",, .---. _

N! t ::I;

v,v,,...':?,

...---"i:,,,......',.. i

'"'

;( ,",'i ,""

",,._,.-:::'.,,"

O.OO4

-- °

1,15

z/D

' I ' I '

I I I I I

.....::.......... _ _ ::, ..,'_

,/.-.'.--:'--., ....._

t" t"t "'r'''" "; ;':_'"'"

It_/ .;;I_t;

O.OO4

-- .5 ' I ' 0.5

FIG"A S(c)

39
'ACE NOZZLE' MODEL

EXPERIMENTAL

FACILITY

NOZZLE WITH

POINTED PLUG

TAB DISCS

6 4% 12_4 % 24_4 % 6_1% 24_1%

FIG. B1

40
SCHEMATIC OF NOZZLE ASSEMBLY

_-1.78 POINTED PLUG

F-.-_] ,l ......

--_ =--,.................... _:_.,_

FIG. B2(a)

NASA/TM_2001-210674 41
SCHEMATIC OF ROUNDED PLUGS

_0.46 3.07 _

d

Interchangeable part

with different porosity

Diameter = 1.57

FIG. B2(b)

NASA/TM--2001-210674 42
LASER SHEET

FLOW VISUALIZATION

POINTED PLUG

Mj = 1.63

NO TAB

24 1%

61%

FIG. B3(a)

NASA/TM--2001-210674 43
LASER SHEET

FLOW VISUALIZATION

POINTED PLUG

Mj = 1.63

24 4%

12 4%

FIG. B3(b)

NASA/TM--2001-210674 44
LASER SHEET FLOW VISUALIZATION

FOR INDICATED PLUG & TAB CONFIGS., Mj -

1.63

SOLID PLUG 5.5%H PLUG

FIG. B4

NASA/TM--2001-210674 45
SCHLIEREN PHOTOGRAPHS

FOR INDICATED PLUG & TAB CONFIGS., Mj = 1.63

SOLID PLUG 5.5%H PLUG

FIG. ns(a)

NASA/TM--2001-210674 46
SCHLIEREN PHOTOGRAPHS

FOR SOLID PLUG, Mj ffi 1.63

FIG. BS(b)

NASA/TM--2001-210674 47
0

Io

K I I I I o

1

k o c_ _ _. o

m I I I I I

,_ I} • • • l)

a a a _

,0

0 0 0 0 0

do66o

0 0 0 0 0 0

H

J ,°

0 _

I

H

0

0

I

d

I I 1 I 1 I I

e_

\

NASA/TM--2001-210674 48
=i

C_

=

_ _, _- _.

l © il ©

1 k

m

0 t_ _

_ 0 0 0

I I I

k

0 _ _ LO

I I I

,I:I il II II

a II a

,0

a 0 0 0

•,_ O_ O_ O_

_D _D qD

0_

0 0 0 0 ,.4 O_ ,_ I

It _ I ._,

. _ _ / r.,_ o

_.ooo ._

I I I I 1 I I I

I ' I ' I l I

8 _ _ 8

V

\

NASA/TM--2001-210674 49
l',,

6 •

_ ._ '_ ,_ _

I _ _._ _

0 II II

I I

•",,__"o '_ _

I I I I

_l I} I} I} •

o o

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I
REPORT DOCUMENTATION PAGE FormApproved

OMB No. 0704-0188

Public reporting burden for this collection of information is estimaled 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, Adington, 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)

4. TITLE AND SUBTITLE

2. REPORTMarchDATE 2001

Effect of Delta Tabs on Free Jets from Complex Nozzles

6. AUTHOR(S)

K.B.M.Q. Zaman

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

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135-3191

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

National Aeronautics and Space Administration

Washington, DC 20546-0001

3. REPORT TYPE AND Technical DATES Memorandum COVERED

5. FUNDING NUMBERS

WU-708-90-43-00

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-12601

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--2000-210674

11. SUPPLEMENTARY NOTES

Responsible person, K.B.M.Q. Zaman, organization code 2600, 216-433-5888.

12s. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category: 02 Distribution: Nonstandard

Available electronically at ht _tp://_ltrs.aTc.na.sa.gov/GLTR$

This laublication is available from the NASA Center for AeroSpace Information. 301-621-0390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum200 words)

Effects of 'delta-tabs' on the mixing and noise characteristics of two model-scale nozzles have been investigated

experimentally. The two models are (1) an eight-lobed nozzle simulating the primary flow of a mixer-ejector

configuration considered for the HSCT program, (2) an axisymmetric nozzle with a centerbody simulating the 'ACE'

configuration also considered for the HSCT program. Details of the flow-field for model (1) are explored, while

primarily the noise-field is explored for model (2). Effects of different tab configurations are documented.

14. SUBJECT TERMS

Mixing; Noise; Tabs; Lobed nozzle

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

18. SECURITY CLASSIFICATION

OF THIS PAGE

Unclassified

NSN 7540-01-280-5500

19. SECURITY CLASSIFICATION

OF ABSTRACT

Unclassified

15. NUMBER OF PAGES

69

16. PRICE CODE

A94

20. LIMITATION OF ABSTRACT

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102