NASA

Technical

Paper

3099

1991

National Aeronautics and

Space Administration

Office of Management

Scientific and Technical

Information Program

Effects of Yaw Angle

and Reynolds Number

on Rectangular-Box

Cavities at Subsonic

and Transonic Speeds

E. B. Plentovich,

Julio Chu,

and M. B. Tracy

Langley Research Center

Hampton, Virginia

Summary

An experimental investigation was conducted to

determine the effect of Reynolds number (separate

from boundary-layer thickness) and the effect of yaw

angle on the pressure distribution in a rectangularbox cavity. The cavity was tested at Mach numbers

from 0.20 to 0.90, Reynolds numbers from 2 x 106

to 100 x 106 ft -1, and yaw angles of 0° and 15 ° .

Cavities were tested with length-to-depth ratios I/h

of 4.4, 6.7, 12.67, and 20.0. Fluctuating- and staticpressure data on the model walls were obtained and a

complete tabulation of the mean static-pressure data

is presented. (The static-pressure data are analyzed

in this report.) The cavity model was mounted

in the sidewall of the Langley 0.3-Meter Transonic

Cryogenic Tunnel. The thickness of the sidewall

boundary layer entering the cavity was measured

with a pitot pressure rake and the tabulated values

are provided. Over the rang e of Reynolds numbers

tested, the ratio of boundary-layer thickness to cavity

depth was approximately constant. There was no

significant effect of Reynolds number on the staticpressure distributions. The effect of yaw on the

cavity pressure distribution was most pronounced

when the flow field was of the open type at 0° yaw.

In such cases the flow field became transitional when

the cavity was positioned at 15 ° yaw. However, if

the flow field at 0 ° yaw was transitional or closed, the

effect of 15 ° yaw on the pressure distribution was very

minimal. This test also showed that the types of flow

field observed for given ranges of l/h at supersonic

conditions would occur for different ranges of I/h at

subsonic and transonic conditions.

Introduction

The flow field within a rectangular-box cavity has

been studied for many years both experimentally and

computationally (refs. 1 to 13). Much of the research

has been directed at developing an understanding of

the flow fields that exist within cavities for the purpose of creating an environment from which stores,

internal to an aircraft, can be released. Most of

the studies have been conducted at relatively low

R%vnolds numbers based on cavity length (R I <_ 15 x

10") and with the cavity aligned with the free-stream

flow direction (0 ° yaw). Therefore, an investigation

was conducted to determine the effect of Reynolds

number at a nearly constant boundary-layer thickness and the effect of yaw on cavity flow fields at

subsonic and transonic speeds. A rectangularbox cavity model was mounted in the sidewall of

the Langley 013-Meter Transonic Cryogenic Tunnel

(TCT). Fluctuating- and static-pressure data on the

model were obtained at Mach numbers from 0.20

to 0.90, Reynolds numbers from 2 × 106 to 100 x

106 ft -1 (Rt from 1.9 x 106 to 94.0 x 106), and yaw

angles of 0° and 15 °. The static-pressure data are

analyzed in this report and the fluctuating-pressure

data are discussed in reference 14.

qoo

cavity depth, in.

cavity length, in.

free-stream Mach number

index in power law for velocity profile

(u zl/n)

p measured surface static pressure, psf

poc free-stream static pressure, psf

pt,oc free-stream total pressure, psf

qoc free-stream dynamic pressure, psf

R free-stream unit Reynolds number,

ft-1

R l Reynolds number based on cavity

length

Tt,_ free-stream total temperature, K

u velocity, ft/sec

Ubl/Uoo ratio of local velocity in boundary

layer to free-stream velocity

w cavity width, in.

x distance in streamwise direction (see

fig. 6), in.

y distance in spanwise direction (see

fig. 6), in.

z distance normal to tunnel sidewall, in.

5 boundary-layer thickness, in.

5* boundary-layer displacement thickness,

in.

0 boundary-layer momentum thickness,

in.

¢ yaw angle, deg

Experimental Methods

Wind Tunnel Description

The tests were conducted in the two-dimensional

adaptive wall test section of the Langley 0.3-Meter

Transonic Cryogenic Tunnel (TCT). A sketch of the

tunnel is presented in figure 1. The 0.3-m TCT

Symbols

Cp pressure coefficient,

h

1

M_

n
is a fan-driven,cryogenicpressuretunnelthat uses

gaseousnitrogenasa testmedium.It is capableof

operatingat stagnationtemperaturesfromapproximately80K to 327K andat stagnationpressures

from 1.2atm to 6.0atm. Thefanspeedisvariable

sothat the Machnumberof theemptytest section

canbevariedcontinuouslyfromabout0.20to 0.95.

Thiscombinationof test conditionsprovidesa test

envelopeofReynoldsnumbersupto about100× 106

basedonamodellengthof 1ft. Additionaldetailsof

thetunnelandits rangeof operationmaybefound

in references15and16.

A sketchshowingdetailsof theflowregionin the

adaptivewall test sectionis presentedin figure2,

andfigure3 isaphotographof thetestsection.The

testsectionis 13in. by 13in. in crosssectionat the

entrance.All fourwallsaresolid. Thesidewallsare

rigid, whereasthetop andbottomwallsareflexible

andmoveable.Theflexibletop andbottomwallsare

computercontrolled,with feedbackprovidedonthe

wall positionandthetop andbottomwall pressure

distribution. The basicobjectiveis to alignthe

test sectionboundarieswith the modelstreamlines

suchthat theflowfieldin thevicinityof themodel

approachesthat whichwouldbe obtainedfor freeair conditions. Whenthe wallsarepositionedin

this way they aresaidto be streamlined. Specific

information on the adaptive wall test section and a

brief description of the strategy used to contour the

walls can be found in reference 17. Specific details

on the contouring strategy for the flexible walls are

given in reference 18.

The 0.3-m TCT model mounting system is designed for two-dimensional models. Typically, the

model is supported between two turntables centered

30.7 in. downstream of the test section entrance.

Models with lengths up to 13 in. can be tested over

an angle-of-attack range of 40 ° . The turntables are

driven by an electric stepper motor that is connected

through a yoke to the perimeter of both turntables.

This arrangement drives both turntables to eliminate

possible model twisting. The angular position of the

turntables and, therefore, the geometric angle of attack of the model are measured with a digital shaft

encoder geared to one turntable.

The 0.3-m TCT has a sidewall boundary-layerremoval system to reduce the boundary-layer thickness (ref. 19). This system uses porous plates

just upstream of the model mounting turntables.

The boundary-layer thickness is not reduced much

with the use of the boundary-layer-removal system

(ref. 19), so it was decided not to use the system

and to replace the porous plates used for boundarylayer removal with solid plates. The location of the

boundary-layer-removal plates relative to the model

position is shown in figure 2.

Model Description

The model tested was a rectangular-box cavity

mounted on the sidewall of the tunnel. The cavity

model was centered on the turntable at about station 0. (See fig. 2.) The position of the turntable

in the test section is shown in figure 2. The model

was fabricated to allow the use of the angle-of-attack

drive for positioning the cavity at yaw angles of

0° and 15 °. Because of the cavity model geometry, the encoder that is generally used to measure

the turntable position could not be utilized for this

test. Therefore, an accelerometer that was calibrated

for 0° and 15° of yaw was used to determine the

turntable and cavity model position. The accelerometer responds to being tilted in the Earth's gravitational field and the output of the accelerometer is a

function of the sine of the angle of inclination. The

accelerometer is capable of resolving angles to within

-t-0.015 ° at 15 ° •

The cavity had a length l of 11.25 in., a width

w of 2.50 in., and a maximum depth h of 2.56 in.

The floor of the cavity could be positioned at various

depths to vary l/h with the length and width fixed.

The 1/h values tested were 4.4 (h = 2.56 in.), 6.7

(h = 1.68 in.), 12.67 (h = 0.89 in.), and 20.0 (h =

0.56 in.). A configuration with the floor mounted

flush with the tunnel sidewall, having no cavity, was

used when the boundary-layer thickness approaching

the cavity was determined. Photographs of the model

are provided in figures 4 and 5. Figure 4 shows the

model prior to tunnel installation, with the floor of

the cavity positioned at a depth of 2.56 in. (I/h =

4.4). Figure 5 shows the model mounted in the tunnel

in the configuration for measuring the boundary layer

(no cavity and the boundary-layer rake installed).

The model was instrumented with 21 staticpressure orifices and 18 flush-mounted fluctuatingpressure transducers, as shown in figure 6. The forward and aft walls of the cavity were instrumented

with only a single fluctuating-pressure transducer located at half the depth of the cavity. Table I provides

the measured positions of the static-pressure ports.

Test Conditions

The model was tested at Mach numbers from

0.20 to 0.90, Reynolds numbers from 2 x 106 to

100 x 106 ft -1, and ¢ = 0° and 15 °. The boundarylayer thickness was nearly constant throughout the

range of Reynolds numbers tested. The model was

tested at a reduced set of conditions and configurations for _p = 15 °. Table II provides a summary
of the nominaltest conditionsandmodelconfigurations.Theflexibletestsectionwallsweresetto a

streamlinedshapefor eachtestcondition.Thisresultedinwalldeflectionsforeachconditionthat were

nogreaterthan±0.04in.for eachwall.

Measurements

Surface static pressures. Because of the large

changes in dynamic pressure in the 0.3-m TCT over

its operational range (a factor of about 75), a highprecision capacitive-type transducer is used for pressure measurements. The electrical outputs from the

transducers are connected to individual signal conditioners. The signal conditioners are autoranging

to keep the electrical output to the data acquisition

system at a high level for all pressure ranges. The

transducers have a maximum range from -100 psi

to 100 psi and have an accuracy of +0.25 percent

of reading from 25 percent of negative full scale to

i00 percent of positive full scale. Additional details

of the 0.3-m TCT pressure instrumentation system

can be found in reference 15.

For the experimental data reported herein, each

orifice was sampled 40 times over a 1-sec period.

These data were then averaged to produce the mean

value for each data point.

Boundary-layer thickness. The ratio of

boundary-layer thickness to cavity depth has been

shown to be an important similarity parameter in

the study of cavity flows (ref. 3). To determine the

boundary-layer thickness at the cavity leading edge

(the tunnel sidewall boundary-layer thickness), the

cavity floor was moved flush with the sidewall (i.e.,

no cavity existed) and the pitot pressure through the

boundary layer was measured with a rake at the cavity leading edge. A drawing of the rake is shown

in figure 7. The boundary-layer rake pressures were

measured with high-precision capacitive-type transducers similar to those used for the surface staticpressure measurements. The reference pressure for

these transducers was the tunnel plenum static pressure. The boundary layer entering the cavity (the

tunnel sidewall boundary layer) has been shown to

be turbulent (ref. 20).

For calculations of Mach number through the

boundary layer, the static pressure at the position

of the rake was needed. (Static pressure through

the boundary layer is assumed to be constant.) The

measurements from the static-pressure ports near

the boundary-layer rake were influenced by the rake

and were therefore in error, so the local wall static

pressure was assumed to be equal to the tunnel static

pressure.

To estimate the boundary-layer thickness _ from

the rake pitot pressure measurements, the data were

reduced using the method described in reference 21.

This method assumes a power-law variation for the

turbulent-boundary-layer velocity profile. A leastsquares fit to the calculated values of Ubl/Uoo is made

to determine the values of _ and n, the power-law

index. Values for displacement thickness _* and

momentum thickness 8 are calculated using numerical integration. For the boundary-layer parameters calculated herein, only the pressure data from

boundary-layer rake tubes within the boundary layer

(Ubl/Uoc < 0.99) were used for the calculations. Because of the nature of the least-squares fit, additional

points outside the boundary layer (within the free

stream) would introduce increasing error into the estimations. The calculated boundary-layer parameters are presented in table III and a plot of the

boundary-layer thickness over the range of Reynolds

numbers tested is shown in figure 8. The pitot pressures through the boundary layer were not measured

at all nominal test conditions, as shown in table II

(I/h = oo), because of tunnel time constraints.

Tabulated data. The cavity pressure measurements were reduced to coefficient form and are presented in tables IV to X. These tables contain the

tunnel test conditions as well as the mean values of

the measured pressures. The pressure data are presented as CPxx, where the xx refers to the orifice

number. (The locations of the orifices are presented

in table I.) Data are presented in order of increasing Mach number and Reynolds number for each

configuration.

Discussion of Results

At supersonic speeds, four types of mean cavity

flow have been defined (refs. 4 and 13) and are

sketched in figure 9. The first type occurs when

the cavity is "deep" (I/h < 10) and is termed open

cavity flow. For open cavity flow, the flow essentially

bridges the cavity and a shear layer is formed over the

cavity. A weak shock can form near the leading edge

of the cavity as a result of the flow being compressed

slightly by the shear layer. The second type of

cavity flow is for "shallow" cavities (I/h > 13) and

is termed closed cavity flow. In closed cavity flow,

the flow separates at the forward face of the cavity,

reattaches at some point along the cavity floor, and

separates again before reaching the rear cavity face.

This creates two distinct separation regions, one

downstream of the forward face and one upstream of

3
therearface.Thethird andfourthmeanflowtypes

(transitional-closedcavity flow and transitional-open

cavity flow) have in the past both been referred to

as transitional flow, that is, where the flow field

changes from closed to open cavity flow. This change

generally occurs for l/h between 10 and 13. The

determination of transitional-closed and transitionalopen flows, as well as determination of open and

closed flows, can best be made by observation of the

pressure distribution in the cavity. Figure 9 provides

a guideline for determining the type of cavity flow,

though it must be recognized that the transonic

characteristics may not be identical to supersonic

characteristics.

Although at supersonic speeds acoustic pressure

fluctuations have been observed in cavities with open

cavity flow, there do not appear to be large variations in the static-pressure measurements. At lower

speeds, large variations in instantaneous pressure

measurements have been observed (refs. 5 and 22).

It is possible to obtain repeatable mean values if a

sufficiently high sample rate is used, but a large variation in amplitude and shape of instantaneous pressure distributions in deep cavities at transonic and

subsonic speeds remains.

Effect of Test Conditions

The data in figure l0 show the effect of tunnel

conditions on the repeatability of pressure distributions on the cavity floor. The 0.3-m TCT can operate over a large temperature and pressure range,

with constant Mach and Reynolds numbers able to

be obtained at many combinations of temperature

and pressure. To address concerns about the effect

of temperature on the dynamic-pressure transducers,

additionM points were included in the test matrix to

repeat a given Mach number and Reynolds number at

different temperatures and pressures. (See table II.)

An attempt was also made to obtain as wide a range

of Reynolds numbers as possible at the same temperature to allow for data comparison in case the

temperature did affect the measurements. The effect of temperature and pressure on M_ and R for

the l/h = 4.4 cavity configuration is shown in figure 10. Figure 10(a) is a comparison of data taken

at M = 0.60 and R --- 30 × 10°at the same temperature and pressure (a repeat point), and figure 10(b)

shows a comparison of data taken at M = 0.60 and

R = 30 × 106 at different temperatures and pressures. For both cases the static-pressure data repeat

well and the tunnel conditions do not significantly

affect the data.

4

Effect of Reynolds Number

The effect of Reynolds number on cavity flow

fields was a principal focus of this test. In most

tests, a change in Reynolds number is usually associated with a change in 5. In this study, however,

the boundary layer approaching the cavity was the

tunnel sidewall boundary layer, which is relatively

thick and insensitive to changes in Reynolds number.

(Fig. 8 provides a plot showing the change in 5 with

Reynolds number.) The value of 5/h changed little

compared with the change in free-stream Reynolds

number (at least a factor of 10), and thus the effects shown are attributed essentially to changes in

Reynolds number.

Figures 11 to 14 show the effect of Reynolds

number at constant Mach numbers for the various

cavity configurations at ¢ = 0 °. As shown in the

plots, there is very little change in the mean Cp

distribution over the range of Reynolds numbers

tested.

The pressure distributions at ¢ = 15 ° are provided in figures 15 to 17. For the majority of the data

there are only slight differences in the measured pressures, and these differences would indicate minimal

effect on flow-field characteristics. The largest influence of Reynolds number on the pressure distribution

is shown in figure 15(a). However, for these data it

appears that the flow field maintains a transitionalclosed characteristic. (Data were not taken for the

I/h = 20.0 cavity configuration at ¢ = 15°.)

Effect of I/h Change

Figures 18 and 19 provide a comparison of cavity pressure measurements for the different cavity

configurations. The cavity length remained fixed at

11.25 in., but depth was varied to generate cavities

with l/h of 4.4, 6.7, 12.67, and 20.0. Because of

the minimal variation in pressure distribution with

Reynolds number this comparison is only shown for

R _ 90 × 106. Figure 18 shows the data for ¢ = 0°,

and as previously discussed, the cavity with l/h = 4.4

has open flow and 1/h = 20.0 has closed flow. The

cavity with I/h = 6.7 was expected to have an open

flow field, but the pressure distribution shows that it

is a transitional-open flow field at M_ = 0.60 and

tends more toward the open type at higher Mach

numbers. Also, a cavity with 1/h = 12.67 would

be expected to be transitional in nature; however,

at M_ = 0.60 the flow field is closed and becomes

transitional-closed at higher Mach numbers.

Figure 19 shows the data for _b = 15 ° and R

90 × 106. Here the flow is transitional-closed for
I/h = 4.4 and 6.7; however, for I/h = 12.67 the flow

is of the closed type.

Effect of Mach Number

The cavity pressure distributions for several Mach

numbers at constant Reynolds numbers are presented

in figures 20 to 23. Each figure shows data for

a separate cavity l/h at ¢ = 0% The data at

Mc¢ = 0.20 have characteristics distinct from those

of the flow fields at higher Mach numbers. These

characteristics were also observed at M_ -- 0.30 in

reference 22 and indicate that the flow-field characteristics at low subsonic speeds are significantly different than the characteristics at higher speeds. As

mentioned previously, in the supersonic speed regime

the 1/h = 6.7 cavity would be expected to be open

and the l/h = 12.67 cavity would be transitional.

However, figure 21 again shows that the l/h = 6.7

cavity is transitional-open at Mc_ = 0.60 and is

changing toward an open type at higher Mach numbers. Figure 22 shows that the l/h = 12.67 configuration changes from a closed cavity flow field to a

transitional-closed cavity flow field as Mach number

is increased. In figure 23 the flow field remains closed

at all Mach numbers tested.

Figures 24 to 26 show the effect of Mach number

on the cavity flow field at ¢ = 15 °. The pressure distributions for I/h = 4.4 and 6.7 give the appearance

of a transitional-closed cavity flow (figs. 24 and 25).

In figure 26, pressure distributions for l/h = 12.67

show that the flow field is closed at Mc¢ = 0.60 and

0.80 and closed on the verge of transitional-closed at

Moc = 0.90.

Effect of Yaw Angle

Figures 27 to 29 present mean pressure distributions for the two yaw angles tested at various

Mach numbers for 1/h -- 4.4, 6.7, and 12.67. (The

I/h = 20.0 cavity configuration was not tested at

¢ = 15°.) Recall that the pressures were measured

along the centerline of the model, so at ¢ = 15° the

pressures are skewed with respect to the free-stream

flow. The effect of yaw angle on a closed or a transitional cavity flow field is minimal. (See figs. 28(c),

28(d), and 29.) However, if the flow field is open at

¢ = 0 °, when the cavity is yawed to 15 ° the flow

field becomes transitional and a much larger change

in the mean pressures is present (figs. 27, 28(a), and

28(5)).

Concluding Remarks

To provide information on the effect of Reynolds

numbers (independent of boundary-layer thickness)

on cavity flow fields at subsonic and transonic speeds,

an experimental study was conducted in the Langley

0.3-Meter Transonic Cryogenic Tunnel. For this

study, cavities with length-to-height ratios I/h of 4.4,

6.7, 12.67, and 20.0 were tested at Mach numbers

from 0.20 to 0.90 and at Reynolds numbers from 2 ×

106 to 100 × l06 ft -1. Static and fluctuating pressures

were measured on the model and the boundary-layer

thickness was measured at the cavity leading edge.

The Reynolds numbers tested (the ratio of boundarylayer thickness to cavity depth was approximately

constant) had no significant effect on the staticpressure distribution. The effect of yaw on the cavity

mean pressure distribution was most pronounced if

the flow field was of the open type at 0° yaw. In

such cases the flow field became transitional when

the cavity was positioned at 15 ° yaw. However, if

the flow field at 0 ° yaw was transitional or closed,

the effect of yaw on the cavity pressure distribution

was very minimal. This test also showed that the

types of flow field observed for given ranges of l/h

at supersonic conditions would occur for different

ranges of l/h at subsonic and transonic conditions.

NASA Langley Research Center

Hampton, VA 23665-5225

April 15, 1991

References

1. Rossiter, J. E.: Wind-Tunnel Experiment on the Flow

Over Rectangular Cavities at Subsonic and Transonic

Speeds. R. & M. No. 3438, British Aeronautical Research

Council, Oct. 1964.

2. Kaufman, Louis G., II; Maciulaitis, Algirdas; and Clark,

Rodney L.: Mach 0.6 to 3.0 Flows Over Rectangular

Cavities. AFWAL-TR-82-3112, U.S. Air Force, May 1983.

(Available from DTIC as AD A134 579.)

3. Charwat, A. F.; Roos, J. N.; Dewey, F. C., Jr.; and Hitz,

J. A.: An Investigation of Separated Flows--Part I: The

Pressure Field. J. Aeronaut. Sci., vol. 28, no. 6, June

1961, pp. 457-470.

4. Stallings, Robert L., Jr.; and Wilcox, Floyd J., Jr.:

Experimental Cavity Pressure Distributions at Supersonic

Speeds. NASA TP-2683, 1987.

5. Dix, Richard E.: On Simulation Techniques for the Separation of Stores From Internal Installations. SAE Tech.

Paper Ser. 871799, Oct. 1987.

6. Heller, Hanno H.; and Bliss, Donald B.: Aerodynamically Induced Pressure Oscillations in Cavities--Physical

Mechanisms and Suppression Concepts. Tech. Rep.

AFFDL-TR-74-133, U.S. Air Force, Feb. 1975.

5
7. Baysal, O.; Srinivasan, S.; and Stallings, L., Jr.: Unsteady

Viscous Calculations of Supersonic Flows Past Deep and

Shallow Three-Dimensional Cavities. AIAA-88-0101, Jan.

1988.

8. Catalano, George D.: Turbulent Flow Over an Embedded

Rectangular Cavity. AFATL-TR-86-73, U.S. Air Force,

Feb. 1987. (Available from DTIC as AD A177 928.)

9. Om, Deepak: Navier-Stokes Simulation for Flow Past an

Open Cavity. AIAA-86-2628, Oct. 1986.

10. Baysal, O.; and Stallings, R. L., Jr.: Computational and

Experimental Investigation of Cavity Flowfields. AIAA87-0114, Jan. 1987.

11. Suhs, N. E.: Computations of Three-Dimensional Cavity

Flow at Subsonic and Supersonic Mach Numbers. AIAA87-1208, June 1987.

12. Srinivasan, S.; Baysal, O.; and Plentovich, E. B.: NavierStokes Calculations of Transonic Flows Past Open and

Transitional Cavities. Advances and Applications in Computational Fluid Dynamics, 0. Baysal, ed., FED-Vol. 66,

American Soc. of Mechanical Engineersl 1988, pp. 169

179.

13. Wilcox, Floyd J., Jr.: Experimental Measurements of

Internal Store Separation Characteristics at Supersonic

Speeds. Store Carriage, Integration and Release, Royal

Aeronautical Soc., 1990, pp. 5.1-5.16.

14. Traey, M. B.; Plentovich, E. B.; Chu, J.; and Shearin,

J.G.: Cavity Acoustics in High Reynolds Number

Transonic Flows. Proceedings 8th JOCG Aireraft/Stores

Compatibility Symposium, Joint Ordnance Commanders

Group, Aircraft/Stores Compatibility Subgroup, and

American Inst. of Aeronautics and Astronautics, Oct.

1990, pp. 25-1-25-28.

i5.

16.

17.

18.

19.

20.

21.

22.

Ladson, Charles L.; and Ray, Edward J.: Evolution,

Calibration, and Ope_tional Characteristics of the TwoDimensional Test Section of the Langley O.3-Meter Transonic Cryogenic Tunnel. NASA TP-2749, 1987.

Rallo, Rosemary A.; Dress, David A.; and Siegle, Henry

J.A.: Operating Envelope Charts for the Langley 0.3-

Meter Transonic Cryogenic Wind Tunnel. NASA

TM-89008, 1986.

Mineck, Raymond E.: Hardware and Operating Features

of the Adaptive Wall Test Section for the Langley 0.3-

Meter Transonic Cryogenic Tunnel. NASA TM-4114,

1989.

Wolf, Stephen W. D.; and Goodyer, Michael J.: Predictive

Wall Adjustment Strategy for Two-Dimensional Flexible

Walled Adaptive Wind Tunnel--A Detailed Description

of the First One-Step Method. NASA CR-181635, 1988.

Murthy, A. V.: Sidewall Boundary-Layer Measurements

With Upstream Suction in the Langley O.3-Meter 7bansonic Cryogenic Tunnel. NASA CR-4192, 1988.

Murthy, A. V.; Johnson, Charles B.; Ray, Edward J.;

Lawing, Pierce L.; and Thibodeanx, Jerry J.: Studies

of Sidewall Boundary Layer in the Langley O.3-Meter

Transonic Cryogenic Tunnel With and Without Suction.

NASA TP-2096, 1983.

Murthy, A. V.: Calculation of Sidewall Boundary-Layer

Parameters From Rake Measurements for the Langley 0.3-

Meter Transonic Cryogenic Tunnel. NASA CR-178241,

1987.

Plentovich, E. B.: Three-Dimensional Cavity Flow Fields

at Subsonic and Transonic Speeds. NASA TM-4209, 1990.

6
Table I. Locations of Static-Pressure Orifices

Orifice no. x, in. y, in. z, in. Model location

1 0 -h Cavity floor

2

3

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

21

0.350

2.050

2.854

3.658

4.462

5.266

6.071

6.874

7.678

8.482

9.286

10.090

10.894

.800

.800

5.666

10.486

10.486

-.690

-.230

11.800

.5

--°5

.5

.5

--.5

.5

.5

.5

Tunnel sidewall, forward of cavity

Tunnel sidewall, forward of cavity

Tunnel sidewall, aft of cavity

7
Table II. Nominal Test Matrix

M(x)

0.20

.20

.20

.6O

.60

.60

.60

.60

.60

.60

.80

.80

.80

.80

.80

.80

.80

.80

.90

.90

.90

.90

.90

.90

R, ft -1

2 x 106

I0

30

4

10

10

30

30

80

90

5

10

10

30

30

80

90

100

I0

30

30

80

90

100

Pt,oo, psi

22.5

26.0

76.0

19.0

46.0

21.0

64.0

30.0

77.0

86.0

18.0

37.0

21.0

62.0

26.0

64.0

72.0

85.0

19.5

68.0

22.5

60.0

67.0

75.0

¢ = 0° and I/h of--

Tt,_c, K 4.4 6.7 12.67 20.0

310 x x

105 x x

105 x x

oo 1

¢ = 15 ° and 1/h of 1

4.4 6.7 12.67

x ×

x X

× ×

200 x x x x x x x x

200 x x x x x

105 x x x x x x x x

105 x x x x x x x x

105 x x x x x x x x

105 x x x x x x x

1Cavity ceiling flush with sidewall for boundary-layer measurements.

310 x x x x x x x x

310 x x x x x x

200 x x x x x x x x

200 x x x x x

105 x x x x x x x x

105 x x x x x x x x

105 x x x x x x x x

105 x x x x x x x x

320 x x x x x x x x

320 x x x x x

180 x x x x x x x x

180 x x x x x

105 x x x x x x x x

105 x x x x x x x x

105 x x x x x x x x
Table III. Boundary-Layer Parameters

Moo R, ft -1 Pt,oc, psi Tt,oc, K 6, in. 6*, in. 0, in.

0.60

0.80

0.90

5 x 106

10

i0

30

30

80

85

6

10

l0

30

30

80

90

100

13

30

80

90

100

20

42

21

63

33

86

88

21

300

299

180

180

115

113

110

300

0.578

.518

.512

.475

.512

.464

.469

.513

0.074

.067

.067

.058

.063

.055

.057

.073

35

23

66

27

71

8O

87

27

26

69

77

86

300

210

209

112

112

112

111

210

114

114

114

114

.544

.505

.489

.494

.475

.463

.437

.511

.491

.463

.456

.458

.071

.070

.064

.066

.059

.060

.059

.073

.064

.062

.060

.061

0.052

.047

.047

.042

.045

.049

.041

.047

.047

.046

.043

.043

.039

.039

.038

.O46

.041

.039

.038

.039

9
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ORIGINAL PAGE

BLACK AND WHITE PHOTOGI_APH

Top flexible wall

d rive rods

Top flexible wall

a

View port

Model mounting

block

Turntable

O

Bottom flexible

wall

L-87:659

Figure 3. Model mounting system of 13-in. by 13-in. adaptive wall test section.

26
ORtG)NAL P ",'3 E

BLACK AND WHITE FHOTOG_APH

L-90-11841

Figure 4. High Reynolds number cavity model prior to tunnel installation.

Boundary-layer rake Flow

Figure 5. Model installed in tunnel.

L-90-11842

27
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m

m

m

m

m

mmmm

mm

m

mm

m

m

mm

m

1.20

0

_-- 1.00 ----_

l

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1.0(

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Figure 7. Boundary-layer rake. All dimensions are in inches.

.60

.55

.50

.45

I I I !

"400 20 40 60 80

Figure 8.

"B

I

100 x

Variation in boundary-layer thickness with Reynolds number.

10 6

29
÷

0

(a) Open cavity flow; l/h < 10,

÷

°__1-

/

(b) Closed cavity flow; I/h > 13.

Figure 9. Flow-field sketches and typical mean pressure distributions for cavity flow types at subsonic speeds.

3O
Cp oj

(c) Transitional-closed cavity flow; 10 < l/h < 13.

Cp

÷

0

(d) Transitional-open cavity flow; 10 < l/h < 13.

Figure 9. Concluded,

31
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.4

Cp .2

0

-.2

Point

m

o 45

[] 69

m ?

[ I I I I

0 .2 .4 .6 .8 1.0

x/1

(a) Repeat points at constant free-stream conditions; Pt,:¢ = 29 psi; Tt,_c = 105°F.

.6-

.4

Cp .2

0

-.2

Pt,oo Tt,oo

o 63.45 179.6

[] 28.54 104.0

?

__ A"] "D

[ I ] I I

0 .2 .4 .6 .8 1.0

x/l

(b) Repeat points at different free-stream conditions.

Figure 10. Repeatability of mean pressure data at Moc = 0.60 and R = 30 × 106 ft -1 for l/h = 4.4.

32
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,51

Report Documentation Page

Nalional Aeronautics and

Space Adminislrat;on

1. Report No. 2. Government Accession No. 3. Recipient's Catalog No.

NASA TP-3099

4. Title and Subtitle

Effects of Yaw Angle and Reynolds Number on Rectangular-Box

Cavities at Subsonic and Transonic Speeds

7. Author(s)

E. B. Plentovich, Julio Chu, and M. B. Tracy

9. Performing Organization Name and Address

NASA Langley Research Center

Hampton, VA 23665-5225

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, DC 20546-0001

5. Report Date

July 1991

6. Performing Organization Code

8. Performing Organization Report No.

L-16847

10. Work Unit No.

505-68-91-12

11. Contract or Grant No.

13. Type of Report and Period Covered

Technical Paper

14. Sponsoring Agency Code

15. Supplementary Notes

16. Abstract

The effects of Reynolds number and yaw angle on pressure distributions in a rectangular-box

cavity were investigated experimentally. The cavity was tested at Mach numbers from 0.20 to 0.90,

Reynolds numbers from 2 to 100 x 106 ft -1, and yaw angles of 0° and 15 °. Cavities were tested

with length-to-depth ratios I/h of 4.4, 6.7, 12.67, and 20.0. Fluctuating- and static-pressure data

on the model walls were obtained and a tabulation of the mean static-pressure data is presented.

The thickness of the sidewall boundary layer entering the cavity was measured and tabulated

values are provided. The Reynolds numbers tested had no significant effect on the static-pressure

distributions. The effect of yaw on the cavity pressure distribution was most pronounced when

the flow field was open at 0° yaw. In such cases the flow field became transitional when the cavity

was positioned at 15 ° yaw. However, if the flow field at 0 ° yaw was transitional or closed, the

effect of 15° yaw on the pressure distribution was very minimal. The types of flow field observed

for given ranges of I/h at supersonic conditions occurred for different ranges of I/h at subsonic

and transonic conditions.

17. Key Words (Suggested by Author(s))

Cavity flow

Transonic speeds

High Reynolds numbers

Pressure measurements

Turbulent boundary layer

19. Security Classif. (of this report)

Unclassified

20. Security Classif. (of this page)

Unclassified

18. Distribution Statement

Unclassified--Unlimited

NASA-Langley, 1991

Subject Category 02

21. 52N°" of Pages 22. A04Price

NASA FORM 1626 OCT 86

For sale by the National Technical Information Service, Springfield, Virginia 22161-2171
)