NASA / TM-1998-208723

Impingement of Boundary-Reflected

Disturbances Originating at the Nose of a

Body of Revolution in the Langley Research

Center 16-Foot Transonic Tunnel

Richard J. Re and Francis J. Capone

Langley Research Center, Hampton, Virginia

National Aeronautics and

Space Administration

Langley Research Center

Hampton, Virginia 23681-2199

October 1998

https://ntrs.nasa.gov/search.jsp?R=19990009542 2018-05-11T16:54:36+00:00Z
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Summary

An investigation has been conducted in the Langley 16-Foot Transonic

Tunnel to determine boundary-reflected disturbance lengths at low supersonic

Mach numbers in the octagonally shaped test section. A body of revolution that

had a nose designed to produce a bow shock and flow field similar to that about

the nose of a supersonic transport configuration was used. The impingement of

reflected disturbances on the model was determined from static pressures

measured on the surface of the model. Test variables included Mach number

(0.90 to 1.25), model angle of attack (nominally -10 °, 00, and 10°), and model

roll angle.

Results showed that the body, which had a blockage ratio of 0.000575,

experienced interference associated with the reflection of the expansion field

generated by the forward portion of the model at the lower supersonic Mach

numbers. The model lengths that would be free of the bow-shock boundaryreflected disturbance were determined at Mach numbers from 1.03 to 1.08. When

compared to bow-shock reflection lengths previously obtained on cone-cylinder

bodies, it was found that the effect of the large nose initial-half-angle

(approximately 15 °) on reflection-free length was lessened by the continuously

decreasing local body half angle downstream of the nose. The variation of

reflection-free length with Mach number was fairly linear and showed essentially

the same trend as cone-cylinder data. Model angle of attack did not affect the

reflected-disturbance impingement point on the model since the center of rotation

for angle of attack was approximately the midpoint of the model. When the model

was rolled to direct the closely-spaced row of pressure orifices towards a test

section slot no difference in pressure distributions was observed. Mach number

1.05 was set with and without the use of test section plenum suction and it was

found that flow removal from the test section through the slots did not affect

boundary conditions enough to be evident in reflection-free model lengths.

Introduction

There are several factors related to model size

that must be considered if valid data is to be

obtained from a transonic wind tunnel test.

Depending on the particular Mach number range of

interest for a given investigation any or all of three

factors must be considered before a model is built or

a commitment is made to a test plan. The first is

model position in the test section; that is, ensuring

that the model is located completely within the

calibrated portion of the test section at all the Mach

numbers of interest. The second is test section

blockage ratio; that is, ensuring that the model

maximum cross-sectional area/test section area ratio

is small enough so that essentially blockage-free data

can be obtained at the highest subsonic Mach

number planned for the investigation. And the third

concerns model length; that is, ensuring that the

model is of appropriate length to avoid impingement

of wall-reflected disturbances at the lowest

supersonic Mach number of the investigation. The

calibration data necessary to allow decisions to be

made on the appropriateness of a certain size model

for an investigation is unique to each wind tunnel

and is often sparse or incomplete especially

regarding shock reflection length since it depends to

some extent on the details of the model geometry. As

time passes the need for more detail of tunnel

capability in specific areas necessitates additional

calibrations. This need forms the impetus for the

present investigation in which data was obtained on

boundary-reflected disturbance lengths in the

Langley 16-Foot Transonic Tunnel for a sharp-nosed

body of revolution at low supersonic Mach numbers.

The Langley 16-Foot Transonic Tunnel is a

closed-circuit single-return atmospheric wind tunnel

that has a slotted transonic test section with a Mach

number range from 0.1 to 1.3. The test section is

octagonal in shape and has longitudinal slots at the

wall vertices. It is fan driven but also uses test-section

plenum suction for Mach numbers of 1.05 and

above. Because of the large test-section crosssectional area it is possible to test small models (with

less instrumention) to higher subsonic Mach numbers

than smaller tunnels before significant blockage

effects are encountered. In addition, the large test

section diameter permits testing of a given model to

lower supersonic Mach numbers (closer to Mach 1.0)

than smaller slotted tunnels before nose-originated

shocks are reflected by the test section boundary and

impinge as disturbances on the model. As the facility

has been modified and upgraded over the years it has

been recalibrated periodically to validate and
document its expanded capabilities. An overview of

the tunnel developmental history, its capabilities, and

detailed current calibrations are contained in

reference 1. Additional tunnel calibrations are

conducted as necessary for specific tests or as part of

comparison surveys with other transonic facilities.

The present investigation was conducted as part

of a survey of three wind tunnels to determine the

impingement location at low supersonic Mach

numbers of wall-reflected disturbances originating as

shocks at the nose of a body of revolution. The other

two wind tunnels included in the survey were the

Boeing Transonic Wind Tunnel (BTWT) and the

Boeing Supersonic Wind Tunnel (BSWT). The nose

of the body of revolution was designed to have a

geometry that would produce a shock having the

same strength as that of a supersonic transport

configuration for which aerodynamic data was

required in the supersonic speed range below Mach

1.1. The body downstream of the nose was designed

to have nearly zero pressure coefficient along its

length at a Mach number of 1.3. From these tests the

lowest supersonic Mach number could be determined

at which data free of wall-reflected disturbance

impingement could be obtained for the specific

model in each of the three tunnels. Impingement of

wall-reflected disturbances on the body of revolution

was determined from the abrupt changes in pressure

measured by a closely-spaced longitudinal row of

pressure orifices. Pressure distributions on the 64-

inch long body in the 16-Foot Transonic Tunnel were

obtained at Mach numbers from 0.90 to 1.25 and at

angles of attack of about -10% 0% and 10 °. This data

complements boundary-reflected-disturbance length

measurements previously made in the 16-Foot

Transonic Tunnel on bodies of revolution with

conical and ogive noses combined with cylindrical aft

sections (ref. 2).

Symbols

Cp

d

db

M

SRB

X

pressure coefficient

local model diameter, in.

model base diameter, 4.605 in.

Mach number

shock reflection body

distance measured from tip of nose aft, in.

TS tunnel station, ft

model angle of attack (measured in vertical

plane of test section), deg

8 test section wall divergence angle, deg

angular location of orifices on model,

positive clockwise, deg

_m model roll angle, positive clockwise, deg

Wind Tunnel

The investigation was conducted in the

Langley Research Center 16-Foot Transonic

Tunnel which is a single-return atmospheric wind

tunnel with continuous air exchange. The test

section is octagonal in shape with 15.5 feet between

the centerlines of opposite walls (equivalent in area

to a circle 16 feet in diameter) and has axial slots at

the wall vertices. The total width of the eight slots

in the vicinity of the model is approximately 3.7

percent of the test section perimeter. The eight flat

segments that comprise the tunnel wall are remotely

actuated and their divergence angle can be changed

while the tunnel is operating to configure the test

section for its optimum calibrated flow quality at

each Mach number. To obtain Mach numbers of

1.05 and above an auxiliary system provides

suction at the outer wall of the test section plenum

and the air removed is exhausted to the atmosphere.

The tunnel arc-sector sting-support system

pivots in the vertical plane (pivot point at TS 134.0)

in such a manner that the model remains on or near

the test section centerline through the angle of attack

range if the model is installed in the appropriate

longitudinal position. A roll coupling is

incorporated in the sting support portion of the arc

sector so that the model can be rolled about its axis

of symmetry (centerline of the test section at o_ =

0 °) without shutting down the tunnel. Since the

tunnel continually exchanges air with the

atmosphere for cooling (20% per circuit) the flow in

the freestream is subject to the daily atmospheric

variations in ambient conditions. The tunnel has

been calibrated at a number of ambient temperature

and humidity conditions and essentially uniform

centerline Mach number distributions are obtained at

Mach _umbers up to 1.1 by setting the test section

walls at specific divergence angles regardless of

ambient conditions. At Mach numbers above 1.1 it

has been found that uniform centerline Mach

number distributions are obtained at specific wall

divergence angles for specific dew point conditions.

The ranges of wall divergence angle set for the range

of Mach numbers of this investigation are presented

in tabl( 1. Details of the tunnel history, operation,

and flo _vqualities are presented in reference 1.

Model

The 64-inch-long axisymmetric model was

designed to have a nose shape that produces a bow

2
shockof aboutthe samestrengthasthe noseof a

realisticsupersonictransportairplaneconfiguration

of the samescale. Photographsof the body of

revolutioninstalledin the16FootTransonicTunnel

are presentedin figure 1 and its longitudinal

locationin thetestsectionis presentedin figure 2.

Thecontoursof the mid- and aft-bodyportionsof

themodelweredesignedusinganiterativetechnique

to haveafree-airpressurecoefficientof nearlyzero

alongtheirlengthat a Machnumberof 1.3. The

nearlyzerodesignpressurecoefficientdistribution

alongthemodelservesasa convenientreferencefor

the detectionof the abrupt changesin pressure

coefficienthatoccurwhentestsectionwall-reflected

disturbancesimpingeon themodel. Only a limited

amountof model geometricdatais availablefor

publicationbutfrom thatit hasbeendeterminedthat

theinitial half angleof thenosewasapproximately

15° and the longitudinalcurvatureon the forward

portionof themodelresultedin achordalhalf angle

of 5.48° for theforward11.750inchesof themodel.

Themodelhada maximumdiameterat its baseof

4.605inches.Geometricdetailsof themodel,model

sting,andthe forward portionof the sting-adapter

arealsoshownin figure2. The101tubesconnected

to themodelpressureorificesexitedthemodelbase

asabundle(fig.l) andwereroutedaft externallyto

the.pressuresensingmoduleswhichwerein the top

port_onof thetunnelsupportsystem.

The anodizedmodelwasmadeof 2024-T3

aluminum and had a polished surface. A

longitudinalrow of 86 closely-spaced0.02-inch

diameterorificeswereinstalledalongthetop (0 =

0°) of the model(t_m= 0°). Additional pressure

orificeswereinstalledat anglesof t_ = 90 °, 180 °,

and 270 ° at 5 stations aft of the model nose (at x =

10, 20, 30, 45, and 60 inches). The locations of the

pressure orifices aft of the nose when the model was

upright in the test section are shown in table 2.

Instrumentation

Model pressures were measured on four 32-

port electronic pressure-scanning modules housed

in the sting-support portion of the arc sector. The

module transducers had a full-scale differential

pressure range of +15 psi and the manufacturer's

quoted worst case static accuracy was +0.10 percent

of full scale. An additional installation effect on

model pressure measurements results from the

sensitivity of the pressure-scanning modules to

temperature. The modules were housed in the

sting-support portion of the tunnel arc sector which

is within the freestream flow environment. Since

the tunnel cooling is dependent on air exchange

with the atmosphere, freestream temperature is a

function of atmosphereic conditions and will also

vary with Mach number and run time. To

minimize the effect of temperature variations on

pressure-scanning module accuracy known

pressures were supplied to two ports of each

module so that the modules could be recalibrated as

needed while the tunnel was operating. Based on

this approach the uncertainty of the measured

pressure coefficients is estimated to be +0.006 at a

Mach number of 0.90 and +0.005 at a Mach

number of 1.2. Model angle of attack was

measured by an accelerometer in the top of the

model sting support system. A correction of 0.1 °

was made for test section upflow angle based on

measurements made on similar models (ref. 1). No

correction has been made to angle of attack for any

small sting deflections that might have occurred

due to model loads.

Tests

The upright model (_m _--" 00) Was tested in the

Mach number range from 0.90 to 1.25 at an angle

of attack of 0 ° with the closely-spaced row of

pressure orifices facing the centerline of the top wall

of the test section. At Mach numbers from 1.00 to

1.25 the model (at 0m = 0° and 90 °) was also tested

at angles of attack of approximately - 10 ° and 10 °.

At Mach numbers of 1.05, 1.15, and 1.25 the model

was rolled (0m) in 22.5 ° increments to 180 ° to

change the orientation of the closely-spaced row of

pressure orifices relative to the flat wall segments of

the test section periphery. The usual method of

setting a Mach number of 1.05 includes the

application of test section plenum suction to avoid

operation of the fan drive at its maximum power.

However, the tunnel is also calibrated for a Mach

number of 1.05 to be set without plenum suction.

Therefore the availability of these two calibrated

methods of obtaining a Mach number of 1.05 allows

the determination of the effect of plenum suction on

the location of wall-reflected-disturbance

impingement on the model. This was done with the

model rolled 90 ° with the closely-spaced row of

model orifices directed toward the centerline of the

side wall of the test section. It is on this side of the

tunnel (at about 80 °) at which the plenum suction is

applied through a 10 foot diameter opening

(butterfly valve).

Since this investigation was directly related to a

test of a supersonic transport configuration in the

16-Foot Transonic Tunnel and the model nose

geometry represented that of the airplane

configuration it was decided to further replicate the

flow field on the nose of the airplane model by

duplicating the artificially fixed boundary layer

transition location selected for the airplane model.

The transition fixing strip was located 1.0 inch aft of
the tip of the noseand consisted of a 0.005-inch

high row of 0.05-inch diameter disks spaced 0.1-

inch apart around the body. Based on the maximum

diameter of the model, which is at the base, the

blockage ratio of the model in the test section

(maximum model cross-sectional area/test section

cross-sectional area) is 0.000575. This indicates that

subsonic data for this model should be free of

blockage effects up to a Mach number of 0.97 (ref.

1).

Presentation of Results

The results of this investigation are presented

graphically as variations of pressure coefficient with

model length in figures 3 through 7. The data are

generally plotted against the length x so that the

dimensional location of boundary-reflecteddisturbance impingement on the model can be

determined directly. As a reference for the

pressure distributions a dashed outline of the body

shape is included along the x-axis in each plot. An

example of the sensitivity of the pressure

coefficient on either side of the reflected-shock

impingement point on the model to tunnel Mach

number setting at low supersonic Mach numbers is

presented in figure 8.

A comparison of the experimental pressure

coefficient distribution measured along the model

at Mach number 1.25 with that calculated using the

Boeing Company's computational fluid dynamics

program TRANAIR is reproduced from reference 3

as figure 9. The nose shock reflection-free model

lengths determined from the present investigation

combined with the cone- and ogive-cylinder

reflection-free lengths determined from previous

16-Foot Transonic Tunnel investigations (refs. 1

and 2) are presented in figure 10.

Discussion of Results

Pressure Coefficient Distributions at a = 0 °

The pressure coefficient distributions along

and around the body at 0 ° angle of attack are

presented in figure 3. These data indicate that the

model was aligned with the flow and on the tunnel

centerline in that pressures at @= 0% 90 °, 180 °, and

270 ° at five model stations are within the range of

the uncertainty of the pressure measurements. At

Mach numbers from 0.90 to 0.98 the pressure

coefficient distributions indicate that pressure on

the nose recovers to essentially freestream static

pressure (Cp = 0) without overexpansion. However,

at a Mach number of 1.00 there is an

overexpansion on the forward portion of the model

that recovers to stream conditions over a length of

about 4 inches. At supersonic Mach numbers up to

1.20 th,'re appears to be impingement on the model

of compression waves that are a reflection of the

expansion waves that occurred over the forward

portion of the model as the flow recovered to

freestream conditions. This effect was seen in the

data of references 2 and 4 for the expansion flow at

the comers of cone cylinders and is discussed in

reference 4. This will be discussed later in the

section entitled Boundary-Reflected Disturbance

Lengths.

As the test Mach number increased and

approached the body design Mach number of 1.3

(design Cp = 0), the pressure distribution became

flatter ahead of the boundary-reflected nose shock

impingement point, the location of which will be

discussed later. There is a pressure increase near

the base of the body at all Mach numbers below

1.20 that is associated with the flow over the model

base and is perhaps also affected by the proximity

of the sting flare and the bluntness of the stingadapter (see figs. 1 and 2). However, the absence

of flow field visualization data and pressure

measurements in the model base and sting area

prevents making any conclusions on the influence

of the sting flare and the blunt sting-adapter.

Experience has shown that it is good practice in

transonic wind tunnel testing to design the model

installation to minimize the effect of the model

support hardware on the aft portion of the model

whenexer possible. In general terms this usually

consists-, of having a sting-to-model base diameter

ratio of about 0.5 with a constant sting diameter

section approximately 5 model base diameters long

followed by a total sting-flare half-angle of less

than 8° . These criteria are considered to be

reasonable to prevent pressurization of the aft

portion of the model by the flow field of the

supporl system and also of assuring that shocks

created by the support system are well aft of the

base legion of the model. Based on the

aforemmtioned "rule-of-thumb" criteria the sting

support system of the present investigation would

fail to meet the constant sting diameter length

requirement. However there should be no effect on

reflected-shock impingement points until the Mach

numbel is large enough for the impingement point

to approach the model base. More detailed

information on the criteria pertinent to sting

interfer,'nce at the model base is contained in

referen,:es 5 and 6.

In order to obtain pressure distributions with

the closely spaced row of model orifices in other

than the vertical plane, the model was rolled about

its centzrline at 0 ° angle of attack at Mach numbers

from 100 to 1.25 (fig.4). This data indicates that

there was no significant effect of model roll

4
orientationon thepressuredistributionsevenwhen

the row of pressureswas in the plane of the

centerlinesof twooppositetest sectionslots(Ore=

22.5°). Theapparenteffectsof body roll angleon

the pressuredistributionsat someof the lower

supersonicMachnumbersareattributableto small

differencesin Machnumbersettingat someof the

roll angles.Thiseffectof smalldifferencesin free

streamMachnumberis illustratedin figure 4(a)

whereat t_m = 0 ° data was obtained at Mach

numbers of 1.030 and 1.032. If the solid circles

are compared to the open circles, a considerable

difference in pressure distributions is evident.

However if the pressures at t_m = 22.5 ° which were

obtained at a Mach number of 1.033 are compared

to the solid symbols, it can be seen that they are

nearly identical.

A graphical representation of the sensitivity of

reflected-shock impingement location to small

variations in Mach number at low supersonic Mach

numbers is presented in figure 8 for the Mach

number 1.05 data of figure 4(b) using greatly

exaggerated scales. At this nominal Mach number

setting, pressure distributions were obtained at 22.5 °

model roll angle increments between 0 ° and 180 °.

However, as roll angle was changed Mach number

varied slightly within the accepted setting tolerance

range of tunnel parameters. This allowed the

selection of two orifices in the steep compression

and expansion portions of the flowfield

downstream of the impingement point to illustrate

graphically the movement of the impingement

point. As can be seen in figure 8 by the opposite

nearly linear trends of pressures measured at the

two locations, the variation in impingement point

resulting from model roll angle (fig. 4(b)) is caused

by the small variations in Mach number.

Effect of Plenum Suction at Mach Number

1.05

The wind tunnel is calibrated to obtain a Mach

number of 1.05 with and without the use of a test

section plenum suction system. When this system is

used, air is removed from the plenum by a

compressor and there is a flow of test-section

boundary-layer air through the slots into the

plenum to replace air that is being exhausted to the

atmosphere. These two methods of setting a Mach

number of 1.05 were included in the test plan to

determine if test section air removal through the

slots affected the test section boundary-reflecteddisturbance characteristics. Data was obtained at

angles of attack of -9.3 °, 0 °, and 10.0 ° and is shown

in figure 5. There was no discernable effect of

plenum suction at 0 ° angle of attack on the pressure

distribution and the small effects on the boundaryreflected-disturbance impingement point on the

body at angle of attack are once again attributable

to the small differences in Mach number shown in

the figure keys.

Pressure Coefficient Distribution at Angle of

Attack

Angle of attack effects on the body pressure

distributions in the supersonic Mach number range

are shown in figures 6 and 7 for the body oriented

at _m = 0 ° and 90 °. With the closely spaced row of

pressure orifices on the top of the model (fig. 6)

angle of attack affected the pressures as would be

expected. That is, at negative angle of attack

(orifice row on the windward side) the bow shock

affecting the top row of pressures was stronger and

therefore a steeper recovery occurred. When the

closely-spaced row of pressure orifices was on the

side of the model (fig. 7) angle of attack in effect

put them in a cross flow. In this case the difference

between the most forward pressure on the nose and

the flatter pressure distribution on the mid-portion

of the model at angle of attack was about the same

as at 0 ° angle of attack. That is, the shape of the

pressure distribution curves were the same at all 3

angles of attack; but the curves were displaced.

Despite the fact that the tip of the nose moved off

the test section centerline at angle of attack the

impingement point for boundary-reflected

disturbances stayed essentially the same as at 0 °

angle of attack (figs. 6 and 7) because the base of

the model moved off the centerline in the opposite

direction since the center of rotation for angle of

attack was approximately at the midpoint of the

model (fig. 2).

Boundary-Reflected Disturbance Lengths

Most of the pressure distributions of figure 3

for supersonic speeds show the impingement of two

major disturbances on the model. The upstream

disturbance impingement point is believed to be

caused by the reflection from the tunnel wall of

expansion waves generated by the contour of the

forebody. In reference 2 it was concluded from

tunnel wall presssure measurements that expansion

waves from the corner formed by the intersection

of the conical nose with the cylindrical afterbody

were reflected back to the model for models having

a blockage ratio of 0.00098 or greater but not for a

model having a blockage ratio of 0.000062. Since

no tunnel wall pressures were recorded during the

present investigation and the model blockage ratio

was 0.000575 it is not possible to know if the

expansion waves reached the tunnel walls.

However, based on the pressure distributions it

appears likely that the expansion waves were

reflected back to the model as compression waves
(seealsoref.4). Thisexpansionfield is peculiarto

modelgeometrydownstreamof the noseand its

strengthis relatedto modelsize. Therefore,since

the subject of this paper concerns boundary

reflecteddisturbancelengthsfor shocksgenerated

bythemodelnoseonly thoselengthsarediscussed

andpresentedin figure 10.

The pressuredistributions for supersonic

speedsof figure3 wereexaminedto determinethe

x locationof thestartof thepressurerisefrom the

boundary-reflecteddisturbanceimpingementof the

noseshock. Determinationof the locationof this

point below a Mach numberof 1.03 was not

possiblebut becameevidentasMachnumberwas

increasedabove1.03. TheweakstrengthandMach

angleof thenoseshockattheMachnumbersbelow

1.03werenot conduciveto determinationof an

impingement point in the pressure

recovery/expansionwaveareaoccurringoverthe

forward portion of the model. The reflecteddisturbanceimpingementlengths for the noseoriginatedshocksarepresentedin figure 10along

with the relevantdatafrom references1 and 2.

Reflected-disturbancelengthdatafrom the current

investigationendsat aMachnumberof 1.08dueto

thelimitedlengthof themodelandtheinfluenceof

theflow field aroundthe modelbaseand support

systemon the pressuressensedby the last 3

orifices. However,the nearlylinear trendsof this

dataandthetrendsobtainedwithlongerbodiesin

references1 and2 indicatethat,if required,some

extrapolationfor longermodelsat higher Mach

numberscanbedonewithconfidence.

The initial half angleof the model noseis

about 15° with local half angle decreasing

continuouslywith lengthdue to the non-conical

modelshapesuchthatat x = 11.375in. the chordal

half angle(tan-1(d/2)/x)is 5.48°. The trend of

boundary-reflected disturbance length for this

investigation shown in figure 10 is approximately

aligned with the 10 ° half angle cone-cylinder data

of figure 10 (model 1) indicating that the influence

of the initial nose half angle on reflecteddisturbance length is lessened considerably when

the pointed nose is not conical. In utilizing the data

of figure 10 for a planned model installation, it

should be remembered that the wing and tail spans

of a model, and model displacement from the

centerline at angle of attack, need to be considered

since the boundary-reflected disturbances may

impinge on model components and still be

downstream of the fuselage base at the centerline of

the tunnel. The effect of model geometry can be

estimated by making a scaled layout of the model

in the test section and using the tip of the model

nose and the impingement length of figure 10 to

draw lines back to a point on the wall (this neglects

any wall boundary layer effect) at the appropriate

shock or Mach angle to determine whether they

pass close to or over any model components.

Bodies having noses that are not sharp will

have quite different reflected-disturbance

impingement lengths in that the nose shock will be

stronger. It will stand ahead of (i.e. be detached

from) the tip of the nose to higher Mach numbers,

and will not bend over to the appropriate Mach

angle -mtil further from the centerline of the

tunnel. In all cases the impingement lengths for

non-pointed bodies will be shorter and will be

greatly dependent on nose/forebody shape. Pre-test

estimation of non-pointed-body reflection-free

lengths for model sizing and test matrix planning

purposes is more uncertain since most tunnels have

very limited calibration data bases for this type of

model geometry. Some early work on the

estimation of detached shock-wave stand-off

distances from the noses of bodies is reported in

references 7 through 9. Unless it is clearly evident

from estimates that there is no possibility of shock

impingement within the planned test matrix,

physical verification of the location of the

reflected-disturbance impingement point by means

of pressure data or flow field visualization is

recommended.

Concluding Remarks

An investigation has been conducted in the

Langley 16-Foot Transonic Tunnel to determine

boundary-reflected-disturbance lengths at low

supersonic Mach numbers. A pressure

instrurrented body of revolution that had a nose

designed to produce a bow shock and flow field

similar to that about the nose of a supersonic

transport configuration was used so that the lowest

reflection-free supersonic Mach number could be

determined. Test variables included Mach number

(0.90 to 1.25), model angle of attack (nominally

-10% 0% andl0°), and model roll angle.

Results showed that the body, which had a

blocka_ e ratio of 0.000575, experienced

interfermce associated with the reflection of the

expansion field generated by the forward portion

of the model at the lower supersonic Mach

numbers. The model lengths that would be free of

the bow-shock boundary-reflected disturbance were

determined at Mach numbers from 1.03 to 1.08.

When compared to bow-shock reflection lengths

previously obtained on cone-cylinder bodies, it was

found that the effect of the large nose initial-halfangle (approximately 15 °) on reflection-free length

was les;ened by the continuously decreasing local

body half angle downstream of the nose. The

variation of reflection-free length with Mach

6
number was fairly linear and showed essentially the

same trend as cone-cylinder data. Model angle of

attack did not affect the reflected-disturbance

impingement point on the model since the center of

rotation for angle of attack was approximately the

midpoint of the model. When the model was rolled

to direct the closely-spaced row of pressure orifices

towards a test section slot no difference in pressure

distributions was observed. Mach number 1.05 was

set with and without the use of test section plenum

suction and it was found that flow removal from the

test section through the slots did not affect

boundary conditions enough to be evident in

reflection-free model lengths.

8. Moeckel, W. E.: Approximate Method for

Predicting Form and Location of Detached

Shock Waves Ahead of Plane or Axially

Symmetric Bodies. NACA TN 1921, 1949.

9. Heberle, Juergen W.; Wood, George P.; and

Gooderum, Paul B.: Data on Shape and

Location of Detached Shock Waves on Cones

and Spheres. NACA TN 2000, 1950.

NASA Langley Research Center

Hampton,VA 23681-2199

October 1998

References

1. Capone, Francis J.; Bangert, Linda S.; Asbury,

Scott C.; Mills, Charles T.; and Bare, E. Ann:

The NASA Langley 16-Foot Transonic Tunnel.

Historical Overview, Facility Description,

Calibration, Flow Characteristics, and Test

Capabilities. NASA TP-3521, 1995.

2. Capone, Francis J.; and Coates, Edward M., Jr.:

Determination of Boundary-ReflectedDisturbance Lengths in the Langley 16-Foot

Transonic Tunnel. NASA TN D-4153, 1967.

3. Haglund, George T.; and Heinrich, Mark R.:

Shock Reflection Probe for Wind Tunnel

Calibration and Investigation of Wall-Reflected

Shocks. AIAA-94-1935, 1994.

4. Estabrooks, Bruce B.: Wall-Interference Effects

on Axisymmetric Bodies in Transonic Wind

Tunnels With Perforated Test Sections. AEDCTR-59-12, ASTIA Doc. No: AD-216698, U.S.

Air Force, June 1959.

5. Cahn, Maurice S.: An Experimental

Investigation of Sting-Support Effects on Drag

and a Comparison with Jet Effects. NACA Rep.

1353, 1958.

6. Kennedy, T. L.: An Evaluation of Wind Tunnel

Test Techniques for Aircraft Nozzle Afterbody

Testing at Transonic Mach Numbers. AEDCTR-80-8, U.S. Air Force, 1980.

7. Laitone, Edmund V.; and Pardee, Otway O'M.:

Location of Detached Shock Wave in Front of

a Body Moving at Supersonic Speeds. NACA

RM A7B 10, 1947.
Table 1. Test Section Wall Divergence Angles Set for the Data Presented

M 5, deg

0.90 4.9

0.92 6.2

0.94 8.0

0.96 10.3

0.98 13.2

1.00 14.2

1.02 9.6

1.03 8.6

1.04 7.3

1.05 6.0

1.06 5.0

1.08 4.0

1.10 1.0

1.15 17.8 or 18.7

1.20 7.6 or 8.7

1.25 14.0 or 15.5
Table 2. Shock Reflection Body Pressure Orifice Locations

$, de{] (_,de{]

x, in. 0 90 180 270 x, in. 0 90 180 270

4.0 • 25.5 •

4.5 • 26.0 •

5.0 • 26.5 °

5.5 • 27.0 •

6.0 • 27.5 •

6.5 ° 28.0 •

7.0 • 28.5 •

8.5 • 29.0 •

8.0 • 29.5 •

8.5 • 30.0 ....

9.0 • 31.0 •

9.5 • 32.0 •

10.0 .... 33.0 •

10.5 ° 34.0 •

11.0 • 35.0 •

11.5 • 36.0 •

12.0 ° 37.0 •

12.5 • 38.0 •

13.0 • 39.0 °

13.5 • 40.0 •

14.0 • 41.0 °

14.5 ° 42.0 °

15.0 • 43.0 •

15.5 ° 44.0 °

16.0 • 45.0 ....

16.5 • 46.0 °

17.0 ° 47.0 •

17.5 • 48.0 •

18.0 • 49.0 •

18.5 ° 50.0 •

19.0 ° 51.0 °

19.5 • 52.0 •

20.0 .... 53.0 •

20.5 ° 54.0 •

21.0 ° 55.0 •

21.5 • 56.0 °

22.0 • 57.0 •

22.5 • 58.0 •

23.0 ° 59.0 °

23.5 • 60.0 ....

24.0 ° 61.0 °

24.5 ° 62.0 °

25.0 • 63.0 •
10

!:iiiii

L-93-13127

(a) Model and arc sector support system.

Figure 1. Photographs showing the body of revolution, model support system and the test

section of the 16-Foot Transonic Tunnel.
0

°_

0

o_

n_

0

11
O"J ',_"1.0

,.ooo

"--0 0_I '_" f.,D

•"a"o c_ e0 ._-

ooo

,-: ,._oo

•--0 I'_ laO0

X0 'l--0 "_"

','- CO QD

0

CO

I--',-

OO

o

l

0

IIIIIIII

I

00

000

XTI

0•

I1) lel- i- t _

if)

OI-.-

e- ._.

°__0

e- ,._ 0 _ 0

_o'_

e,-

_ e-

°__

0

0 _-

--•0X00o

0 r"

{=. *_c0 0E 0 0 .-- 0 •

_am

ai IJ_

12
co 0

--:F - ̧ _ .... ! ......

M = 0.901

i i

_i ii I _!i ; i;, i_i ii

c;

co

0

0

-7,7-$i--7--i

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

_:_i. Iii : iii 2: IL Ii}i :i} i }]_: i

;1:1 .........::! ̧ ̧ ̧ ̧11:"i:::11:: /ii

M= 0.938 :,

i k

O, deg

o 0

• 90.0

,_ 180.0

• 270.0

co

co

0

-.05

_-i ......... i ̧i ̧:i ̧i ̧ ̧1 ̧ ̧ ̧ii  ̧I-;':I  ̧I:1:1111 ' ii

-:; ' _;; ; : ; , M=0.959

....xi ̧171:;17 .... :: .... :x 111:

:i iiiiiii_i ii , ! ii;

.10

.05

0

--_ i l-_i.:_il ., ! .........................

-.05 q:iiiiiii:i iii ii:iiii i i:ii,:i:i ............. -- -- i

0 10 20 30 40 50 60 70

x, inches

(a) M = 0.90 to 0.98.

Figure 3. Pressure distributions on the body at subsonic and supersonic

Mach numbers for _z = 0 ° and Om = 0°.

13
cp

.15_ _, _ .......

=!! ! ! I!ii !ii i}!lii:;:::::: :i::: :i_,:: ii

- _ _ l i i _ i i .... i I _ t

- i } !LI i. i ........... !i ...... . !i_Ii! i. ii iii : I ii:: i

.10 - iii i ,;

Z :7 : ii ii I i :; I ;ii iii ; _: ..... ; ; , 117 : ;i, iil;i:: ;:;:: ::i:;: ;7

.O5

co

0

: : : , :: M=I.O00

_.I'_ ----,u,..,, '

x: :::'x::::::::x:xT'::::: x::

.15

.10

=i: ii , it ii . i; ..... !; .iii/iii : ililiiill ii_llii : ii

.O5

0

¢, deg

0

90.0

180.0

270.0

c;

0

-.05 M, =1.021 i i_ : ii:

.15

•05 -: O:::;:::'::i::: :' ::!i:: !;:E::!!:::::i:i :::i : ::];i:; :::::i:!:;::'::::_.........:;::

0

= :; : : :i::l::: :x'::x :::l :: :x :::x:i :: :x,:x ::::::::::x::x::: .....: :x  ̧ ̧

: ;; ,:: M = 1.032 ' ':' ..........

-.05 -311i' ii:i i l:i ;:i:iii; i;l::A::i;:i:_ ............ -- 'i, il:iii i

0 10 20 30 '4'0' 50 60 70

x, inches

(b) M = 1.00 to 1.03.

Figure 3. Continued.

14
.15

.10

Cp .05

0

.10

Cp .05

- ii::i :; 7; ? I _i ; i

M = 1.041

I t

0

{P,deg

0

90.0

180.0

270.0

0

.10

.... ' M = 1.049

t I

i  ̧ _, _l i  ̧ :i

Cp .05

}

M= 1.060

,,,,,,_, _,_,t_,t_ ,,,,,_, ................. ililiiii_ _:i:iiii: ii:iiiii:il -.05

0 10 20 30 40 50 60 70

x, inches

(c) M = 1.04 to 1.06.

Figure 3. Continued.

15
cp

.15

.10

.05

0

.10

.05

=:i: : :: : : I::i : :::::::i::4 ̧:: ̧::;:::;:;; : I::: :;:::: :;: I : i: ::; :: :i::l:::;::i........i .... :::i;:::::: ;::::

"ii ! _i [_ii i-ii_[ ii i! i_ i  ̧ ii iii .......

..................... : ..................... , ....... ! ̧ ̧ ̧, , -w, r........ : .............................. : " "ii' ": ....... r...... i ! i

ill::! i:::_i!_

-;: : :,:;;:;:;, ...... : :::!:: :::

M = 1.080 .........

I , ,:

t

"i! iii iii ii ; i_:l_i, :,i _ _ k i i ii i k , i ,i

: ; :; :; ::::: i::1::::; ......... ::7;

0

_, deg

0

90.0

180.0

270.0

cp

0

.10

.05

-,i .... , ,, ,_. ,..............

................ M _ | .099

: I I ! : ] :: I I

Zl i; :i i:: : ; i:ii:, i:ii:: :/:::i :; i:i::i ?, :i[:::i: :i........... '."I::i:i::: ::i i::

_iil ii[ iiii i:iii t iiiijiillill ilill Iii: i ̧ii:ifliiiii!liiiiiii i !i ̧

........... I " _........ i ....... i .... _ : • ' " : ..... : .... t _ ........... ! .......... : : ............. i ...........

0

- 05 -qlli*iiiiti:i:i;:ii .....

: : : M= 1.151 r:i::i::;:i:

liitlilll

"--0 10 20 30 40 50 60 70

x, inches

(d) M = 1.08 to 1.15.

Figure 3. Continued.

16
.15

.10

.05

;: ;YIII_,;I ;i:: :i ;i, : i;ii :;ii ,:: ;i ; :;i: k ̧ O

¢, deg

0

90.0

180.0

270.0

cp

0

-.05

.15

.10

.05

,, M= 1.199

J i

_:!:i i!!il!ill i:i!li:!! !!i!ii i!!! ii!!l i:: !!! :: :::::i!!!11 ::: i::

....... '' "''1 "' ....... I ..... ,_ E _, ......... , .... I ''_ ,"

_ .ilk

"+iii k,, [i. . i iiii .[ 'IIZI ii[i'iii ..... i '"

-': ..... ..... M=1.251 .. . I :

-.05 qiii,,iii'iiii_iiilriiii,ii: ................. iiiiiliiiiiiiiiiiii_iiii

0 10 20 30 40 50 60 70

x, inches

(e) M = 1.20 and 1.25.

Figure 3. Concluded.

17
co

113

0

M

o 1.000

o 1.002

t_ 0.996

_m, deg

0.0

22.5

90.0

-.05

.10

.05

0

-.05

.10

.05

0

_,' iiiiiiiiii_iiii!iil iiiiiiiiii!iiiiiiiiiii '_i i1

-.05

0 10 20 30 40 50 60 70

x, inches

M _m, deg

o 1.020 0.0

[] 1.021 22.5

1.020 90.0

M Cm, deg

o 1.030 0.0

• 1.032 0.0

[] 1.033 22.5

_- 1.030 90.0

(a) M = 1.00 to 1.03.

Figure 4. Effect of model roll angle on pressure distributions measured by

closely spaced longitudinal row of orifices, a = 0 °.

18
cp

cp

cp

.10

.05

0

.05

0

.05

x, inches

(b) M = 1.04 to 1.06.

Figure 4. Continued.

O

[]

<>

A

r_

r_

13

O

¢

M

1.041

1.040

1.042

M

1.049

1.051

1.051

1.052

1.051

1.052

1.051

1.050

1.049

_m, deg

0.0

22.5

90.0

_m, deg

0.0

22.5

45.0

67.5

90.0

112.5

135.0

157.5

180.0

M _m, deg

o 1.060 0.0

[] 1.059 22.5

tx 1.057 90.0

19
cp

.10

.05

0

M _m, deg

o 1.080 0.0

n 1.077 22.5

_- 1.080 90.0

cp

cp

.O5

.o5 -|k _li I i

0

5dlfflllll lllllIll! llllllll[ llllllllr lllllllll |11[1111! llllllllI -" 10 20 30 40 50 60 70

x, inches

M _m, deg

o 1.099 0.0

[] 1.098 22.5

tx 1.100 90.0

M _m, deg

o 1.151 0.0

[] 1.152 22.5

<> 1.151 45.0

lX 1.152 67.5

t,, 1.151 90.0

r-, 1.151 112.5

c3 1.151 135.0

o 1.150 157.5

(c) M = 1.08 to 1.15.

Figure 4. Continued,

2O
co

co

.10

.05

0

0

10 20 30 40 50

x, inches

60 70

M Om, deg

o 1.199 0.0

[] 1.200 22.5

v. 1.201 90.0

M era, deg

o 1.250 0.0

[] 1.250 22.5

<> 1.249 45.0

ZX 1.249 67.5

tx 1.249 90.0

r', 1.248 112.5

a 1.249 135.0

o 1.251 157.5

0 1.250 180.0

(d) M = 1.20 and 1.25.

Figure 4. Concluded.

21
c;

c;

c;

.05

0

-.05

-.10

.10

.05

0

-.05

.05 __- I. : :l_:! ] " i: ' '

±

" 0 10 20 30 40 50 60 70

x, inches

0

0

0

M a, deg

1.050 -9.29 I

1.051 -9.30 ]

I

1

J

M a, deg

1.051 0.01

1.051 0.00

M a, deg

1.049 10.00

1.050 10.00

Figure 5. Pressure distributions on the body at a Mach number of 1.05

with test section plenum suction on and c_ff,Cm = 90°. (Open

symbols indicate that test section plenum suction is being applied

and solid symbols indicate plenum suction is off.)

22
.20

.15

M t_, deg

o 1.003 -9.31

[] 1.000 0.00

o 1.000 10.00

cp

.10

.05

0

cp

.15

.10

.05

M tz, deg

o 1.016 -9.31

[] 1.020 0.00

<> 1.019 10.01

0

-05 ,;;,,i:,;:,{i:,;,,,:,,,{;:,,,,,:,,:,{;:,i,,,,i:,l;,;:,,;:,i,

• 0 10 20 30 40 50 60 70

x, inches

(a) M = 1.00 and 1.02.

Figure 6. Effect of angle of attack on body pressure distributions measured

with the closely spaced longitudinal row of orifices on top of the model, Om = 0°"

23
.20

cp

co

.15

.10

-.05 : .....

.20

_-k i] I I i : :: I :_ ..... : :

.15

.10

.05

M a, deg

o 1.029 -9.31

[] 1.030 0.00

<> 1.028 9.99

M a, deg

o 1.041 -9.29

o 1.041 0.00

o 1.042 10.00

0

"050 10 20 30 40 50 60 70

x, inches

(b) M = 1.03 and 1.04.

Figure 6. Continued.

24
co

cp

.20

0

i! _i!ilill !i ii!ilk!!i!!ii!ii!!iiii!i!iiiiii!i i_! ii!iiiiii!iil

!iil ̧ i ] !!i! !ii

.20 = : : : : : k ! : : : : : k ̧

-.0 -_,ii,iiiiliiiiiiii,liiii,iiiiliiii,iiiilii_i_iiiil,i,_,,,iiliii_,,,,i

50 l0 20 30 40 50 60 70

x, inches

M a, dog

o 1.051 -9.30

[] 1.049 0.01

<> 1.049 10.00

M a, deg

o 1.060 -9.60

[] 1.060 0.00

o 1.059 10.00

(c) M = 1.05 and 1.06.

Figure 6. Continued.

25
cp

.20

.15

.10

05 _ --

M a, dog

o 1.080 -9.60

[] 1.080 -0.01

o 1.079 10.00

0

.15

.10

.05

M a, deg

o 1.098 -9.32

[] 1.099 0.00

<> 1.099 10.00

0

- 05 ii_,i;"ii;l'ii;i'ii;

• 0 10 20 30 40 50 60 70

x, inches

(d) M = 1.08 and 1.10.

Figure 6. Continued.

26
cp

.20

.15

.10

.05

M a, deg

o 1.149 -9.61

[] 1.151 0.01

0 1.151 9.99

0

co

.15

.10

.05

0

-.05

I

i i i i!!iiii

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

-]iiiiiiii'i,iil,iii'iii,iiiiiliiiiiii,i'iiiiiii:i:i'iiiiiiiii'_iiiiiiii

0 10 20 30 40 50 60 70

x, inches

M a, deg

o 1.199 -9.61

[] 1.199 -0.01

<> 1.199 10.02

(e) M = 1.15 and 1.20.

Figure 6. Continued.

27
co

M a, dog

1.250 -9.60

1.250 0.00

1.250 10.00

(f) M= 1.25.

Figure 6. Concluded.

28
c;

c;

.10

.05

0

-.05

-.10

.05

0

-.05

M a, deg

o 1.003 -9.30

[] 0.996 0.01

<> 0.999 10.00

M a, dog

o 1.019 -9.30

[] 1.020 0.01

(> 1.020 10.00

-.10 ........._-:-

"150 10 20 30 40 50 60 70

x, inches

(a) M = 1.00 and 1.02.

Figure 7. Effect of angle of attack on body pressure distributions measured with

the closely spaced longitudinal row of orifices on the side of the model, Cm= 90°"

29
.lO I _! ....

.05

0

-.05

M or, deg

o 1.032 -9.30

[] 1.030 0.0l

<> 1.031 10.00

.10 ....

!ii!i!i_-i! i_iii M a, deg

o 1.041 -9.29

[] 1.042 0.00 I

<> 1.041 10.00 i

1

x, inches

(b) M = 1.03 and 1.04.

Figure 7. Continued.

30
co

.lO

.O5

0

-.05

M a, deg

o 1.050 -9.28

[] 1.051 0.01

<> 1.049 10.00

-.10 ........

.lO

-:: :! I: ̧ ::!:1 :: :: kl!: ..... :!: :!: ! ! i:1 : !

05

0

" 0 10 20 30 40 50 60 70

x, inches

M a, deg

o 1.059 -9.28

[] 1.057 0.00

<> 1.059 10.01

(c) M = 1.05 and 1.06.

Figure 7. Continued.

31
M _, deg

o 1.080 -9.29

[] 1.080 0.00

<> 1.083 10.00

0

-.05

.10 ................. ...........

.05 _i

0

-.05

M o_, deg

o 1.103 -9.29

o 1.100 0.01

<> 1.102 10.00

-.10

_,i,,,i,;,,i;,,;it,i;;iiiiii,,,;iil;ili,il;i;i,ii,liiii,,ilil;

-'150 10 20 30 40 50 60 70

x, inches

(d) M = 1.08 and 1.19.

Figure 7. Continued.

32
.10

.05

M a, deg

o 1.148 -9.29

o 1.151 0.01

1.149 10.00

co

0

-.05

co

15

.10

0

x, inches

M a, deg

o 1.201 -9.29

[] 1.201 0.00

1.203 10.01

(e) M = 1.15 and 1.20.

Figure 7. Continued.

33
.10

M a, dog

o 1.249 -9.29

[] 1.249 0.00

<> 1.249 10.01

cp

x, inches

(f) M = 1.25.

Figure 7. Concluded.

34
.06

x, in.

© 40.0

[] 47.0

.05

.04

.03

.02

.01

\

I I I I I

1.049 1.050

/

I I I I I I I I I I I

1.051 1.052 1.053

M

Figure 8. Variation of pressure coefficient with small variations in

Mach number at two locations on the model in the vicinity of

Mach number 1.05.

35
Data

.25

o Exp

TRANAIR

.20

.15

.10

.05

0

-.05

--

,-._,___:'___ _ _ ---__ _,___- __

111111111 IIII1|111 IIIIIIIII IIIIIIIII I|11|111|

0 10 20 30 40

I

IIIIIIIII IIIIIIIII

50 60 70

x, inch_s

Figure 9. Comparison between measured and calculated

pressure distributions at cL= 0° and M = 1.25.

36
O

[]

<>

A

&

C

9

Model Length, in. Nose half angle, deg Location Reference

1 96.0 10 Centerline 2

2 96.0 20 Centerline 2

3 96.0 Ogive Centerline 2

5 60.0 10 Centerline 2

9 100.3 14 Centerline 2

21 72.0 14 +22 in. off centerline Unpublished

22 96.0 14 +22 in. off centerline Unpublished

SRB 64.0 Centerline Current

100

_=-

t_

ttO

O

tt'_

L..

°--

a

9O

80

70

60

5O

40

30

20

10

0

1.00

Shock interference //_ 13 ,/"

/ .

:1//

Shock free

-I I I I I I I I I I I I I I I I I I I I I I I t I I I I I I I I t I

.04 1.08 1.12 1.16

! I

Illllll li till

1.20 1.24 1.28

M

Figure 10. Boundary-reflected-disturbance lengths measured in the

Langley 16-Foot Transonic Tunnel for nose originated shocks.

37
REPORT DOCUMENTATION PAGE Form Approved OMB NO, 0704-0188

Publlc,reporttng I_ for..this collection of information !s estimated to ave*age 1 hour pet..r_:qponse, indudircg the time for revl_vtng tnstruct_ns, searching existing data sources,

.ga:olenng ar_. malnta,inlng _ .0eta need, _ co .m_/ng ..a_:l.revlewtng the colleen of information_ Send com.iTle_S re_jard.k'lgthis burden estimate or any other aspect of this

co_,De, c[,c_ oT minim^a_on,_nc_ua.!ngsug_sz.._ons rot mouong mls our_, to Wash=ngmn Headquarters _ervmes, Directorate _r Information Operations and Reports 1215 Jefferson

uaws _Kjnway, _une 1_,u4, _qlngton, VA 22L:_2-.4_UZ, and to the uno_ of Management and Budget, Paperwork Reduction Project (07040188), Wasltngton, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE I 3. REPORT TYPE AND DATES COVERED

October 1998 I Technical Memorandum

4. TITLE AND SU6TiTLE 5. FUNDING NUMBERS

Impingement of Boundary-Refiected Disturbances Originating at the Nose of

a Body of Revolution in the Langley Research Center 16-Foot Transonic WU 537-07-20-21

Tunnel

6. AUTHOR(S)

Richard J. Re and Francis J. Capone

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

NASA Langley Research Center

Hampton, VA 23681-2199

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

National Aeronautics and Space Administration

Washington, DC 20546-0001

11. SUPPLEMENTARY NOTES

8. PERFORMING ORGANIZATION

REPORT NUMBER

L-17755

10. SPONSORING/MONffORING

AGENCY REPORT NUMBER

NASA/TM-1998-208723

12a. DiSTRIBUTION/AVAILABILITY STATEMENT • 12b. DISTRIBUTION CODE

Unclassified-Unlimited

Subject Category 02 Distribution: Nonstandard

Availability: NASA CASI (301) 621-0390

13. ABSTRACT (Maximum 200 words)

An investigation has been conducted in the Langley 16-Foot Trans )nic Tunnel to determine boundary-reflected

disturbance lengths at low supersonic Mach numbers in the octagonally shaped test section. A body of

revolution that had a nose designed to produce a bow shock and flow field similar to that about the nose of a

supersonic transport configuration was used. The impingement of reflected disturbances on the model was

determined from static pressures measured on the surface of the model. Test variables included Mach number

(0.90 to 1.25), model angle of attack (nominally -10 °, 0°, and 10°), and model roll angle.

14. SUBJECT TERMS

Shock reflection; Boundary reflected disturbances;

Transonic wind tunnel test section; 16-Foot Transonic Tunnel

17. SBCURITY CI-PSSIRCATION 18. SECUHilV C_SIRCATION 19. SECURITY CLASSIRCATION

OF REPORT OF THIS PAGE OF AB.CrRACT

Unclassified Unclassified Unclassified

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