February 2010 

NASA/TM–2010-216197 

Unified Application of Vapor Screen Flow 

Visualization and Pressure Sensitive Paint 

Measurement Techniques to Vortex- and 

Shock Wave-Dominated Flow Fields 

Gary E. Erickson 

Langley Research Center, Hampton, Virginia 
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February 2010 

NASA/TM–2010-216197 

Unified Application of Vapor Screen Flow 

Visualization and Pressure Sensitive Paint 

Measurement Techniques to Vortex- and 

Shock Wave-Dominated Flow Fields 

Gary E. Erickson 

Langley Research Center, Hampton, Virginia 
Available from: 

NASA Center for AeroSpace Information 

7115 Standard Drive 

Hanover, MD 21076-1320 

443-757-5802 

Acknowledgments 

The authors would like to thank Dr. Blair B. McLachlan and Dr. James H. Bell of NASA 

Ames Research Center for their contributions that led to the establishment of a pressuresensitive paint system at the NASA LaRC 7- by 10-Foot High Speed Tunnel, 8-Foot 

Transonic Pressure Tunnel, and Unitary Plan Wind Tunnel. Their assistance in the 

definition, installation, and operation of the PSP system and the provision of the 

PAINTCP image processing software are greatly appreciated and respectfully 

acknowledged. 

The authors would also like to thank the PSP team headed by Mr. Clifford J. Obara at 

NASA LaRC, who provided technical assistance throughout the PSP testing in the cited 

facilities. In particular, the contributions of Mr. Amir J. Jagharghi, Mr. B. Scott Sealey, 

Mr. Cecil G. Burkett, Mr. Bradley D. Leighty, and Michael T. Carmine are 

acknowledged with gratitude. 

Special thanks are extended to the research and technical staff assigned to the Unitary 

Plan Wind Tunnel and 8-Foot Transonic Pressure Tunnel who supported the testing 

discussed in this report. Their professionalism and support are appreciated and 

respectfully acknowledged. 
v

Table of Contents 

Table of Contents .......................................................................................................................................... v

Abstract ......................................................................................................................................................... 1

Introduction ................................................................................................................................................... 1

Nomenclature ................................................................................................................................................ 2

Flow Visualization and Measurement Techniques ....................................................................................... 2

Laser Vapor Screen ................................................................................................................................... 2

Pressure Sensitive Paint ............................................................................................................................ 3

Wind Tunnel Facilities, Test Conditions, and Data Corrections .................................................................. 3

NASA LaRC UPWT Facility Description and Test Conditions ............................................................... 3

NASA LaRC 8-Foot TPT Facility Description and Test Conditions ........................................................ 4

Data Corrections in the NASA LaRC UPWT and 8-Foot TPT ................................................................ 4

Model Descriptions ....................................................................................................................................... 4

Generic 65-Degree Cropped Delta Wing-LEX ......................................................................................... 4

Generic Double-Delta Wing ..................................................................................................................... 5

Faceted Missile ......................................................................................................................................... 5

Discussion of Results .................................................................................................................................... 6

Generic 65-Degree Cropped Delta Wing-LEX ......................................................................................... 6

Generic Double Delta Wing .................................................................................................................... 10

Faceted Missile Model ............................................................................................................................ 13

Concluding Remarks ................................................................................................................................... 15

References ................................................................................................................................................... 15

Figures..........................................................................................................................................................17 
vi

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1

Abstract 

Laser vapor screen (LVS) flow visualization and pressure sensitive paint 

(PSP) techniques were applied in a unified approach to wind tunnel 

testing of slender wing and missile configurations dominated by vortex 

flows and shock waves at subsonic, transonic, and supersonic speeds. 

The off-surface cross-flow patterns using the LVS technique were 

combined with global PSP surface static pressure mappings to 

characterize the leading-edge vortices and shock waves that coexist and 

interact at high angles of attack (). The synthesis of LVS and PSP 

techniques was also effective in identifying the significant effects of 

passive surface porosity and the presence of vertical tail surfaces on the 

flow topologies. An overview is given of LVS and PSP applications in 

selected experiments on small-scale models of generic slender wing and 

missile configurations in the NASA Langley Research Center (NASA 

LaRC) Unitary Plan Wind Tunnel (UPWT) and 8-Foot Transonic 

Pressure Tunnel (8-Foot TPT). 

Introduction 

The understanding and control of leadingedge vortex flows, shock waves, mutual 

interactions of vortices and shock waves, and 

flow-field interactions with airframe 

components is a continuing challenge for 

designers of modern military and commercial 

aircraft and missile configurations. Wind tunnel 

testing is an important method of visualizing and 

quantifying the aerodynamic characteristics of 

advanced aerospace vehicle configurations at 

subsonic, transonic, and supersonic speeds. A 

multitude of test techniques exist to examine the 

on-surface and off-surface flow field 

characteristics, many of which feature optical 

methods that are well-suited to reveal the 

complexities of vortex- and shock wavedominated flow fields in the subsonic through 

supersonic speed regimes. The LVS flow 

visualization method (reference 1) was used in 

selected experiments described in this report to 

illuminate the cross-flow patterns about slender 

wing and missile models. The off-surface flow 

visualization results were combined with global 

surface static pressure response maps using the 

PSP technique (reference 1). Previous 

experience (references 2-4) indicates that the 

LVS and PSP techniques are complementary, 

since each technique affords new insights or 

corroborative findings regarding the physics of 

the vortex flows and shock waves that form on 

slender wing and slender body configurations. 

The application of the two techniques is 

mutually exclusive, however, because of 

competing light sources and locations for the 

LVS and PSP optical components, different 

model surface preparations, and wind tunnel test 

condition requirements. Consequently, the two 

techniques were applied in separate phases of 

the selected experiments, which required 

appropriate planning and efficient use of wind 

tunnel facility resources. Representative results 

obtained on three selected configurations in 

separate experiments are presented in this report. 

One experiment was conducted in the NASA 

LaRC UPWT using a 65-degree cropped delta 

wing model with leading-edge extension (LEX) 

and centerline and twin, wing-mounted vertical 

tails. The LEX was tested with and without 

passive surface porosity for vortex control. Two 

experiments were performed in the NASA LaRC 

8-Foot TPT using a double-delta wing model 

featuring a 76o

/40o strake-wing planform and a 

faceted missile model with chine-like cross 

sections. Detailed discussions of the results 

obtained in the UPWT and 8-Foot TPT testing 

of the cropped delta wing-LEX, double delta 

wing, and faceted missile models are provided in 

references 2, 3, and 4, respectively. 
2

Nomenclature 

B.L. model butt line, inches 

BMC balance moment center 

c wing centerline chord, inches 

Cp static pressure coefficient 

ESP electronically-scanned pressure 

LaRC Langley Research Center 

LEX leading-edge extension 

LVS laser vapor screen 

M free-stream Mach number 

MRC moment reference center 

M.S. model station, inches 

NASA National Aeronautics and Space 

Administration 

PSP pressure-sensitive paint 

Re Reynolds number, millions per 

foot 

s semispan distance, inches 

TPT Transonic Pressure Tunnel 

UPWT Unitary Plan Wind Tunnel 

W.L. water line, inches 

x distance along wing centerline 

chord, inches 

y distance along local semispan 

measured from the wing 

centerline, inches 

 angle of attack, degrees 

 angle of sideslip, degrees 

Flow Visualization and 

Measurement Techniques 

Laser Vapor Screen 

The vapor screen method of flow 

visualization has been used in wind tunnel 

testing for several decades to visualize vortices, 

vortex sheets, lines of flow separation and 

reattachment, and shock waves at subsonic, 

transonic, and supersonic speeds. Water is 

injected into the wind tunnel circuit in a 

controlled manner and in sufficient quantity to 

promote condensation of water vapor in the test 

section. A laser is often used to produce an 

intense sheet of light that is projected into the 

test section typically in a plane perpendicular to 

the longitudinal axis of the tunnel or to the body 

axis of the model. At supersonic speeds, the 

temperature drop from the expansion in the 

supersonic nozzle causes the water vapor to 

condense into a thin, uniformly-distributed fog. 

The distribution of condensed water vapor and, 

consequently, the amount of scattered light 

within the plane of the light sheet, is affected by 

the flow disturbances created by the model. 

This phenomenon permits the observation and 

documentation of vortex cross sections, for 

example, at high angles of attack as shown in the 

vapor screen image in figure 1 (reference 5). 

Condensation first appears in the free stream at 

supersonic speeds, so that vortex flows appear as 

dark regions in the absence of scattered light 

surrounded by a light background. The change 

in flow density through oblique shock waves 

results in a similar change in fog density so that 

shock positions and shapes are often clearly 

defined. Separated flows such as wakes, vortex 

feeding sheets, and vortex core regions appear as 

dark or transparent, since condensate does not 

appear to be convected across the shear lines. 

At subsonic and transonic speeds, the condensed 

water vapor generally appears near the central 

region of the vortices, so the vortex cross 

sections appear as light regions within a darker 

background. A combination of the two lightscattering patterns often occurs at transonic 

speeds. LVS results are considered qualitative, 

although relative positions, sizes, and shapes of 

vortices and shock waves can be extracted from 

the vapor screen images, as desired. The 

appearance of condensation in the test section 

will affect the free-stream flow characteristics 

and the quantitative measurements of the model 

surface pressures and forces and moments 

(reference 1). Condensation at supersonic 

speeds is accompanied by a stagnation pressure 

loss and a decrease in the Mach number at the 

condensation shock (reference 1). 

Interpretations of the vapor screen images are 

still valid, however, and effective correlations 

can be made with the trends observed in 

quantitative model measurements such as 

surface static pressure distributions and sixcomponent forces and moments. Details of the 

LVS systems that were used in the NASA LaRC 

UPWT and 8-Foot TPT are described in 

references 2-4. The light sheet was directed 

through optical-quality windows in the side of 

the test section for UPWT testing. The position 
3

of the light sheet was fixed and was oriented 

perpendicular to the free-stream. The model 

was traversed forward and aft through the fixed 

light sheet at a given angle of attack using the 

axial translation capability of the model support 

system. For testing in the 8-Foot TPT, the light 

sheet-generating optical package was positioned 

in the test section ceiling. The light sheet passed 

through an optical-quality window centered in 

the ceiling. The optics package included a 

remotely-controlled rotating mirror that directed 

the light sheet in an arc along the length of the 

fixed model at a given angle of attack. 

Consequently, the light sheet was typically nonorthogonal to the model surface. The LVS 

images that are presented in this paper are 

digitized frames acquired from a personal 

computer video capture board connected to a 

miniature camera installed inside the test section 

on the model support system (reference 1). 

Pressure Sensitive Paint 

PSP is a global optical surface static pressure 

measurement technique that is based on the 

oxygen-quenching characteristics of certain 

luminescent materials. The emitted light 

intensity varies inversely with the local oxygen 

partial pressure and, therefore, the air pressure, 

since oxygen is a fixed mole fraction of air. The 

key elements of a PSP system include 

photoluminescent material in the form of a paint 

applied to the test article, illumination source to 

excite the paint, imaging device such as a 

scientific-grade digital camera to document the 

paint in the excited state, and an image 

acquisition and processing system. An in-situ 

global calibration method was used in the NASA 

LaRC wind tunnels to calibrate the PSP. 

Specifically, pressure measurements were 

obtained at several discrete pressure tap 

locations on the model surfaces using an 

electronically-scanned pressure (ESP) 

measurement system, and the PSP intensity was 

calibrated from the pressure tap data at spatially 

corresponding locations. Every pixel in the 

painted portion of a PSP digital image is 

effectively a pressure tap, so the pressure 

distributions can be resolved to much greater 

detail in all applicable regions of the model. 

Similar to surface oil flow visualization, colorcoded images depicting the calibrated PSP 

intensity field response (that is, the surface 

pressure response) provide pertinent information 

on the topology of the vortex flows and shock 

waves that induce the corresponding pressure 

distributions. Figure 2 (reference 1) shows a 

color-coded surface pressure response obtained 

on a slender wing at a supersonic Mach number. 

The potential intrusive effects of the PSP on the 

aerodynamic characteristics can be controlled by 

minimizing the paint thickness and surface 

roughness. An assessment of PSP intrusiveness 

on the aerodynamic flow about selected models 

at subsonic through supersonic speeds is 

provided in reference 6. Details of the PSP 

systems used in the NASA LaRC UPWT and 

8-Foot TPT are also described in references 2-4. 

Wind Tunnel Facilities, Test 

Conditions, and Data Corrections 

NASA LaRC UPWT Facility Description 

and Test Conditions 

The NASA LaRC UPWT is a continuousflow, variable-pressure supersonic wind tunnel. 

The tunnel contains two test sections which are 

approximately 4 feet square and 7 feet long. 

Each test section encompasses only part of the 

Mach number range of the tunnel. The nozzle 

ahead of each test section consists of an 

asymmetric sliding block which allows 

continuous Mach number variation during tunnel 

operation from 1.5 to 2.86 in the low Mach 

number test section and 2.3 to 4.63 in the high 

Mach number test section. A complete 

description of the facility is contained in 

reference 7. 

The experiment using the 65-degree 

cropped delta wing-LEX model was performed 

in UPWT Test Section 1 at free-stream Mach 

numbers of 1.6 and 2.1, a Reynolds number of 

2.0 million per foot, and a stagnation 

temperature of 120 degrees Fahrenheit (deg F). 

Optical access to the test section is through a 

series of windows in the side walls. This 

required the model to be rolled to a wingsvertical position for the PSP phase of the 

experiment. Mechanical operation of the model 
4

support system in the yaw plane constrained the 

angle of attack to a maximum of approximately 

8 degrees. The model was tested in the upright 

orientation for the LVS phase of the testing, 

which afforded an increased range of angle of 

attack using the standard pitch mechanism. 

NASA LaRC 8-Foot TPT Facility 

Description and Test Conditions 

The NASA LaRC 8-Foot TPT was 

designed for operation as a continuous-flow, 

closed-return, variable-pressure wind tunnel 

with control capability to independently vary 

Mach number, stagnation pressure, stagnation 

temperature, and humidity. The test section was 

square with corner fillets and a cross-sectional 

area approximately equivalent to that of an 

8-foot diameter circle. The top and bottom walls 

of the test section were axially slotted to permit 

a continuous variation of the test section Mach 

number from 0.2 to 1.2. A detailed description 

of the 8-Foot TPT is provided in reference 8. 

The testing of the double delta wing model 

was conducted at free-stream Mach numbers of 

0.50, 0.85, and 1.20, a Reynolds number of 2.0 

million per foot, and a stagnation temperature of 

80 deg F. The testing of the faceted missile 

model was also performed in the 8-Foot TPT at 

free-stream Mach numbers of 0.90 and 1.20, a 

Reynolds number of 3.0 million per foot, and a 

stagnation temperature of 120 deg F. Optical 

access through a window positioned along the 

test section ceiling allowed testing of the models 

in the upright orientation for the LVS and PSP 

phases of the experiments. This provided the 

full range of angle of attack using the arc sector 

model support system. 

Data Corrections in the NASA LaRC 

UPWT and 8-Foot TPT 

The 6-component force and moment data 

obtained in the NASA LaRC UPWT and 

8-Foot TPT were corrected for test section flow 

angularity. The axial force data were also 

corrected to the condition of free stream static 

pressure acting on the body chamber area. The 

model base areas were negligible, hence, no base 

pressure measurements were required. The 

model angle of attack was corrected for sting 

and balance deflections due to aerodynamic 

loads. Blockage and buoyancy corrections were 

not applied. 

Model Descriptions 

Generic 65-Degree Cropped Delta WingLEX 

A generic fighter model featuring a 65-

degree cropped delta wing with sharp leading 

edges was used in the UPWT testing. The 

model was designed and fabricated for surface 

pressure and force and moment testing in 

subsonic, transonic, and supersonic wind 

tunnels. The wing had an NACA 64A005 airfoil 

section from the 40-percent chord station to the 

trailing edge. A sharp, symmetric leading edge 

was obtained by fairing a biconvex circular-arc 

section into the NACA profile from the 

40-percent chord station to the wing leading 

edge. The wing was mounted in a high position 

on a fuselage that served as a housing for 

balance and pressure instrumentation. The 

wing-fuselage configuration represented a 

baseline configuration for the International 

Vortex Flow Experiment on Euler Code 

Validation in the 1980s, and extensive 

experimental and computational results obtained 

by the international community are summarized 

in reference 9. The model was modified to 

include a wing LEX mounted to an integral strut 

or ‘gooseneck,’ a centerline vertical tail, and 

twin vertical tails on the wings. A flat-plate, 

0.25-inch thick LEX having a 65o

/90o planform 

and symmetrically-beveled leading edges was 

fabricated for this experiment. The exposed area 

of the LEX (left and right sides) was 15 percent 

of the reference wing area. Reference 2 includes 

a more detailed discussion of the model 

geometry. The LEX incorporated a pattern of 

0.05-inch diameter through-holes spaced 0.10-

inch apart on center to provide a total porosity 

level of 14.75 percent relative to the LEX 

exposed area. The objective of this passive 

porosity concept was to control the strength of 

the LEX vortices and their interaction with the 

wing vortex flows for improved high angle-ofattack aerodynamics. The same LEX was tested 

with 0 percent porosity (solid LEX) by applying 
5

sealing tape having 1.8 mil thickness (0.0018 

inches) along the lower surface to cover all of 

the through holes. The model with porous LEX, 

wing, and twin wing-mounted vertical tails is 

shown installed in UPWT Test Section 1 in 

figure 3. A photograph of the model rolled 90 

degrees in preparation for PSP testing is also 

shown in figure 3. Planview and sideview 

sketches of the wing, LEX, fuselage, and tails 

are provided in figure 4. The right wing upper 

surface was instrumented with a total of 45 

pressure orifices distributed in three spanwise 

rows. The pressure rows were located at 30 

percent, 60 percent, and 80 percent of the 

distance, x, along the wing centerline chord, c, 

measured from the apex of the wing (x/c = 0.30, 

0.60, and 0.80). The orifice nondimensional 

semispan location, y/s, is expressed in terms of 

the semispan distance, y, measured from the 

wing centerline divided by the wing local 

semispan, s. Consequently, y/s values of 0.0 and 

1.0 correspond to the wing centerline and the 

right wing leading edge, respectively. The 

pressure orifices were connected to an internal 

ESP module with purge air capability, which 

allowed air to be routed through the orifices 

during the application of the PSP coating. This 

avoided having to mask off the pressure rows 

during the painting process. 

Generic Double-Delta Wing 

The generic double delta wing model 

tested in the 8-Foot TPT was also designed and 

fabricated for surface pressure and force and 

moment testing in subsonic, transonic, and 

supersonic wind tunnel facilities. The 0.375-

inch-thick wing featured a flat upper surface and 

sharp, asymmetrically-beveled leading, side, and 

trailing edges. The leading-edge sweep angles 

of the highly-swept forward section, or strake, 

and the main wing panel were 76o

 and 40o, 

respectively. A fiberglass housing was attached 

to the lower surface of the wing and served as a 

protective cover for the ESP and balance 

instrumentation. A more detailed description of 

the model is provided in reference 3. Low-speed 

wind tunnel and computational fluid dynamics 

results obtained on this configuration prior to the 

8-Foot TPT testing are summarized in 

reference 10. Photographs of the double delta 

wing installed in the 8-Foot TPT test section are 

presented in figure 5 and correspond to the 

unpainted model and the model with PSP 

coating applied. Both photographs show the 

model with modified strake-wing intersections, 

or fillets (reference 3), which are not discussed 

in this report. The planview and sideview 

sketches in figure 6 show the dimensional details 

of the model. The model incorporated a total of 

47 static pressure orifices distributed in three 

spanwise rows on the wing upper surface. The 

three rows were on the left-hand side of the wing 

at nondimensional axial positions of x/c = 0.25, 

0.75, and 0.90, which correspond, respectively, 

to 25 percent, 75 percent, and 90 percent of the 

distance, x, measured from the strake apex along 

the model centerline chord, c. The orifice 

nondimensional semispan location, y/s, is 

defined in the same manner as the 65-degree 

cropped delta wing model. Consequently, y/s

values of 0.0 and -1.0 correspond to the wing 

centerline and the left wing leading edge, 

respectively. The pressure orifices were 

connected to a standard ESP module without 

purge air capability. Consequently, the pressure 

rows were masked off during the PSP coating 

application. These regions appear as unpainted 

strips in the processed PSP images. 

Faceted Missile 

The faceted missile model tested in the 

8-Foot TPT was a 30-percent-scale slender 

lifting body representation of a proposed missile 

concept investigated in a cooperative 

experimental research program involving NASA 

LaRC and the United States Navy's Naval Air 

Systems Command, China Lake, California. 

The model was designed for the dual purpose of 

pressure and force and moment measurements, 

and it was compatible with testing in subsonic, 

transonic, and supersonic wind tunnels. The 

body was faceted and featured chine-like cross 

sections with sharp, symmetric leading edges. A 

more detailed description of the model is 

provided in reference 4. Figure 7 presents 

selected wind tunnel installation photographs of 

the missile model, which include the model 

unpainted and with PSP coating applied to the 

upper surface. Overall dimensions of the missile 

configuration are shown in figure 8. The model 
6

top and side views and cross sections can 

generally be described as diamond shapes. The 

model had an expanding cross section forebody, 

a transitional region, and an aftbody with 

contracting cross section. The development of 

vortex flows from the sharp leading edges was 

expected. It was also anticipated that the flow 

field over the model would exhibit asymmetries 

at sufficiently high angles of attack because of 

the slenderness of the model and the possible 

development of multiple vortices arising from 

the discontinuities in the model planform and 

changes in the cross-sectional shape along the 

length of the model. ESP pressures were 

measured at selected stations by full rings of 

orifices. In terms of the local semispan, pressure 

orifices were placed at the body centerline (0 

percent semispan) and in 5-percent increments 

to 95 percent semispan. The model length was 

36 inches, and the pressure rings were located at 

model stations (M.S.) 9.0, 20.50, 25.50, 30.00, 

and 34.50 inches. This arrangement provided 

one ring on the forebody (M.S. 9.00), one ring 

on the transition region (M.S. 20.40), and three 

rings on the aftbody (M.S. 25.50, 30.00, and 

34.50). The pressure orifices were connected to 

standard ESP modules installed inside the 

model. Similar to the double delta wing model, 

the pressure rows were masked off during the 

application of the PSP coating, and the pressure 

rows appear as unpainted strips in the PSP 

images. The field-of-view that was selected for 

the PSP testing was the model upper surface and 

included the pressure measurement stations at 

M.S. 20.40, 25.50, and 30.00. The ESP 

measurements and LVS flow visualization 

results shown in this report correspond to these 

model stations. 

All models were tested without artificial 

surface roughness for boundary layer transition, 

since primary boundary layer separation was 

assumed to occur at the sharp leading edges at 

the high angles of attack of interest. 

Discussion of Results 

Generic 65-Degree Cropped Delta WingLEX 

Figure 9 presents PSP and LVS results 

obtained on the 65-degree cropped delta wing 

model with solid LEX and porous LEX at 

M = 1.6 and  = 8o

. The false-colored PSP 

images at the top of figure 9 show the wing 

upper surface pressure response. The LEX was 

unpainted and, therefore, is not visible in the 

PSP images, which have undergone registration, 

ratioing, and resection transform (reference 1). 

Image registration marks (small black circles) 

are visible in the PSP images. The model 

fastener holes were initially filled with 

automotive putty, and these holes are also 

visible as larger dark circles through the PSP 

coating. The holes were filled with dental 

plaster prior to repeat applications of the PSP, 

which eliminated the appearance of the fastener 

holes in the processed images. A color bar 

indicates that green, blue, and purple correspond 

to regions of progressively lower pressure (that 

is, higher suction pressures), with purple 

representing the highest suction pressure levels. 

Regions of higher surface pressure (lower 

suction pressures or slightly positive pressures) 

are represented by shades of yellow, red, and 

pink. The LVS images for each configuration 

are shown directly below the corresponding PSP 

image, and they depict the condensation patterns 

in cross sections at approximately 60 percent, 80 

percent, and 100 percent of the distance along 

the wing centerline chord measured from the 

wing apex (x/c = 0.60, 0.80, and 1.00, 

respectively). The three x/c stations are 

identified as red dashed lines in the PSP images, 

and the x/c = 0.60 and x/c = 0.80 stations are 

locations at which discrete spanwise surface 

static pressure measurements were acquired 

along the right wing upper surface using the ESP 

method. The x/c = 1.00 station is at the wing 

trailing edge. The left portion of the LVS image 

at x/c = 1.00 is in the shadow of the centerline 

tail, since the laser light sheet was projected into 

the test section through a window in the righthand wall. The LVS images are scaled relative 

to each other to reflect the growth of the vortices 
7

from the forward to aft portions of the wing. 

The LVS images are also approximately scaled 

relative to the PSP images so that the vortex 

cross sections in the LVS images can be 

projected to the corresponding regions of higher 

suction pressures in the PSP images. Important 

features of the PSP and LVS images are 

identified with numbers, and a descriptive key is 

provided in the figure. The location of the LEXwing junction is also denoted. 

The PSP image obtained with the solid 

LEX in figure 9 reveals subtle pressure 

signatures induced by the LEX vortices and 

denoted by narrow bands of dark green/light 

blue colors extending in a nearly streamwise 

direction along the wing. There is a region of 

higher pressures denoted by yellow, red, and 

pink along the forward region of the wing and 

between the LEX vortex pressure footprints. A 

dominant feature of the PSP image are the 

broader bands of blue and purple color 

extending inboard from the wing leading edges, 

which are associated with the wing vortical 

flows. The footprint of an oblique shock wave 

beginning near the apex of the centerline tail and 

extending outboard across the LEX vortex 

pressure signatures is also apparent. The shock 

is revealed as a narrow, swept band of yellow 

and red colors extending outboard on either side 

of the tail. The LEX and wing vortex pressure 

signatures are in proximity, but there is no 

indication of direct interaction (that is, 

intertwining of the vortices) or instability of the 

vortices. It is noted that vortex instability, or 

breakdown, would not be expected at these test 

conditions (reference 11) in the absence of large 

adverse streamwise pressure gradients caused, 

for example, by high angle of attack, a strong 

shock wave, or the presence of a solid obstacle 

in the path of the vortices. The LVS images 

show two distinct vortex cross sections on either 

side of the wing centerline: a LEX vortex of 

nearly circular cross section positioned along the 

inboard portion of the wing and surrounded by a 

band of water vapor condensate, and a wing 

vortex situated along the outer section of the 

wing which appears as a broader, flatter region 

of low or negligible water vapor condensate. 

The wing vortex ‘feeding sheet’ is 

approximately defined in the LVS images as the 

boundary between the low condensate region 

within the vortex and the nearly uniform 

condensation region outside the vortex. The size 

of the LEX vortex cross section belies the 

relatively small pressure signature in the PSP 

image. The LEX vortex is nearly circular 

because it is a ‘free’ vortex. It detaches from the 

leading edge near the LEX-wing junction and 

does not receive additional leading-edge 

vorticity as it passes over the wing. The 

condensate that is entrained between the LEX 

vortices implies a region of induced downflow 

that correlates with the higher pressure region in 

the PSP image. The flatter wing vortex cross 

section is also consistent with the broader 

signature in the PSP surface pressure mapping. 

The LVS images indicate the wing vortex is 

situated closer to the wing upper surface. In 

addition, the vortex feeding sheet is attached to 

the entire wing leading edge. These LVS flow 

field features correlate with the region of more 

intense blue and purple colors in the PSP image, 

which are indicative of higher vortex-induced 

suction pressures. The LVS images also reveal 

the presence of condensation between the LEX 

and wing vortices, which is a region of vortexinduced flow that reattaches to the wing upper 

surface. This reattached flow region is coded 

typically a green color in the PSP image and is 

situated between the boundaries of the LEX and 

wing vortex pressure footprints. The LVS 

image near the wing trailing edge (x/c = 1.00) 

reveals the presence of the LEX, wing, and wing 

tip vortices and a trace of the centerline vertical 

tail shock. The LVS result at x/c = 1.00 supports 

the inferences from the PSP image, namely, the 

LEX and wing vortices do not directly interact 

and the vortices are stable to the wing trailing 

edge. There is no indication in the LVS image 

that the traversal of the LEX vortex across the 

boundary of the shock emanating from the tail 

causes instability of the vortical flow. The PSP 

image in figure 9 corresponding to the porous 

LEX features a single, broad wing vortex 

pressure signature on each side of the wing 

centerline and an absence of the LEX vortex 

pressure footprints. In addition, regions of 

higher surface pressures along the forward and 

central sections of the wing and on either side of 

the centerline tail are less extensive, which 

suggest a mitigation of the vortex-induced 
8

downflow and a reduction of the shock strength 

at the tail. Certain trends in the PSP image 

correlate with the corresponding LVS images, 

which indicate that LEX porosity shifts the 

dominance from the LEX vortex to the wing 

vortex. The singular feature in the LVS images 

is a broader, thicker wing vortex cross section 

compared to its solid LEX counterpart. There is 

no evidence of the LEX vortices, and the 

induced downflow above the center portion of 

the wing is less discernible. No inferences can 

be made, however, regarding the reduced shock 

strength at the tail that was suggested in the PSP 

images. 

Figure 10 presents expanded views of the 

right-hand portions of the PSP images with solid 

LEX and porous LEX shown previously in 

figure 9. The corresponding LVS flow 

visualization of the right half of the wing at 

x/c = 0.80 is shown directly below the PSP 

images. The x/c = 0.80 station is denoted as a 

red dashed line in the PSP image. The PSP and 

LVS images are scaled relative to the plots of 

the spanwise distributions of the right-hand wing 

upper surface static pressure coefficient, Cp, 

shown at the bottom of the figure. Cp is plotted 

against the nondimensional semispan location, 

y/s. Two distributions are shown in each plot 

corresponding to discrete pressure 

measurements obtained using the ESP system 

and from the PSP image which was calibrated 

via the global in-situ method. The dual vortex 

pressure footprints and vortex cross sections 

with the solid LEX correlate with the two 

pressure signatures in the Cp distributions. One 

might infer from the relative size of the LEX 

vortex in the LVS image that the latter would 

induce a higher suction pressure peak compared 

to the wing vortex. This is not the case, 

however, as the PSP pressure map and Cp

distributions indicate the LEX vortex induces a 

minor suction pressure peak compared to the 

higher suction pressure plateau induced by the 

broader, flatter wing vortex. The Cp distribution 

obtained with the porous LEX is shown at the 

bottom right in figure 10, which captures a 

single suction pressure plateau along 

approximately the outer 60 percent of the wing 

semispan and an inboard region of increasing 

pressure (that is, reduced suction pressure). The 

lateral extent of the suction pressure plateau in 

the presence of the porous LEX correlates with 

the broadened vortex pressure signature and 

vortex cross section in the PSP and LVS images, 

respectively. In addition, the region of wing 

vortex-induced flow reattachment inferred from 

the PSP and LVS images is consistent with the 

pressure rise inboard of the suction pressure 

plateau. 

Many details of the flow about the 

65-degree cropped delta wing can be derived 

from the combined results obtained using the 

PSP and LVS techniques. However, results 

presented so far in figure 9 and figure 10 do not 

clearly identify the effect of flow-through 

porosity on the LEX vortex system. A technique 

often used in wind tunnel flow visualization 

experiments is to perform an angle-of-attack 

‘sweep’ from low-to-high values to observe, for 

example, the development and growth of 

vortices, vortex interactions, and shock waves. 

Frequently, the enlarged scale of the flow at 

higher angles of attack allows one to better 

discern flow-field features of interest that still 

exist but are less visible at lower angles of 

attack. This approach could not be adopted 

using the PSP method in the UPWT testing, 

since the optical access required the model to be 

oriented in a wings-vertical position in the test 

section and the angle of attack was limited by 

mechanical operation of the support system to 

approximately +8 degrees in this orientation. 

LVS flow visualization was performed with the 

model in the standard wings-horizontal 

orientation, and the support system was capable 

of pitching the model to angles of attack 

significantly higher than 8 degrees in this plane 

of motion. Blockage considerations and balance 

instrumentation load limits at M = 1.6 precluded 

angles of attack greater than approximately 12 

degrees. However, these constraints were eased 

at M = 2.1 where the flow-field features were 

topologically similar to M = 1.6 (reference 2). 

Figure 11 shows the LVS cross-flow patterns at 

M = 2.1,  = 16o

, and x/c = 0.80 with the solid 

LEX and porous LEX. At this higher angle of 

attack, the LEX and wing vortices are visible in 

both cases. With the solid LEX, the vortices are 

distinct and separated, and multiple cross-flow 

shock waves develop above, between, and below 
9

the vortices. With the porous LEX, the 

identities of the LEX and wing vortices can still 

be discerned, but the vortices join to produce a 

cross flow pattern that is characteristic of a 

single, broader vortical flow. The LVS image 

also suggests the combining of the vortices 

changes the pattern of cross-flow shocks that 

bound the vortex flows. These results confirm 

that porosity does not suppress the LEX vortex 

flows but, instead, causes a diffusion and 

weakening of the vortices. The phenomena 

observed at this higher angle and Mach number 

are expected to exist, albeit to a smaller scale, at 

the conditions previously shown in figure 9 and 

figure 10. Specifically, the redistribution of the 

PSP surface pressure response, the LVS crossflow patterns, and the static pressure coefficient 

distributions caused by porosity is induced by 

diffuse, weakened LEX vortices that have 

combined with stronger and larger wing vortical 

flows. It is noted that a similar conclusion 

would likely not have evolved with the exclusive 

use of the PSP method. 

Figure 12 shows the PSP surface pressure 

maps and LVS cross-flow images at M = 1.6 and 

 = 8o corresponding to the solid LEX in 

combination with the centerline vertical tail 

(depicted on the left) and with twin, wingmounted vertical tails (depicted on the right). 

The LVS images were acquired at x/c = 0.60, 

0.80, and 1.10, where the latter station is in the 

near wake of the wing. The centerline and twin 

vertical tails are positioned downstream of the 

80 percent wing chord station, and, their 

upstream influence at supersonic speeds is 

limited to disturbances transmitted through the 

wing boundary layer. As a result, the PSP 

surface pressure response and LVS cross-flow 

patterns forward of the tail positions are similar 

for both tail arrangements, and significant 

differences are confined to the local regions 

about the tails and into the near wake. As noted 

previously in figure 9, the signature of a shock 

emanating from the apex region of the centerline 

tail is apparent in the PSP image, and this shock 

intersects the pressure footprints of the LEX 

vortices. The LVS cross-flow pattern at 

x/c = 1.10 reveals the tail shock, which has no 

apparent effect on the stability of the LEX or 

wing vortices. A direct interaction of the wing 

and LEX vortices occurs in the near wake. The 

wake roll-up and wing tip vortex are also 

apparent in the LVS image at this station. The 

PSP and LVS images are more complex in the 

vicinity of the twin vertical tails. The PSP 

surface pressure response shows the LEX vortex 

footprints intersecting the leading edges of the 

wing-mounted tails. Downstream of this 

location, the LEX vortex pressure footprints are 

no longer visible in the PSP image. A strong, 

intersecting shock system is established between 

the twin tails, and a rapid pressure rise across the 

shocks is indicated in the PSP image by the 

transition from a green color ahead of the 

shocks, to yellow near the foot of the shocks 

and, finally, to shades of red and pink 

downstream of the shock fronts. It is noted that 

the PSP color display was adjusted during the 

image processing in an attempt to clearly show 

the important features in the surface pressure 

response. In doing so, the paint intensity 

response in certain small sections of the PSP 

image would sometimes exceed the extremes of 

the available color palette, and these saturated 

regions appear as white or black in the images. 

The PSP response outboard of the twin tails 

reveals a swept band of blue and purple colors, 

which suggests a region of organized flow 

separation (that is, vortex flows). Outboard of 

this region, the higher suction pressures induced 

by the wing leading-edge vortex are manifested 

as a broad band of blue and purple colors. The 

LVS image at x/c = 1.10 enhances the 

interpretation of the PSP image. Interestingly, 

although the PSP image suggests the LEX 

vortices directly impinge on the vertical tails, the 

LVS flow visualization shows that the LEX 

vortices retain their structure and stability into 

the near wake, even as they traverse through the 

higher pressure region established downstream 

of the twin tail shock system. The central region 

of the LEX vortex is shown as a dark area at 

about the tail midspan on the inboard side of 

each vertical tail. The traces of the tail shocks 

are also evident in the LVS cross-flow pattern. 

The LVS image confirms the development of a 

separated flow region containing low or 

negligible water vapor condensate situated 

outboard of the twin tails. The cross flow 

induced by the LEX vortices at the tails, in 

combination with the tail leading-edge sweep 
10

angle of approximately 45 degrees, may be 

sufficient to promote a partial-span vortical flow 

positioned near the junction of the tail and wing 

upper surface. The presence of this separated 

flow region limits the inboard lateral growth and 

movement of the wing leading-edge vortex, 

which is visible adjacent to the wing tip vortex 

in the LVS image at x/c = 1.10. 

A more detailed examination of the flow 

behavior in the vicinity of the right-hand wingmounted vertical tail at M = 1.6 and  = 8o

 is 

provided in figure 13. PSP static pressure 

coefficients were extracted from two chordwise 

rows positioned at nondimensional semispan 

locations of y/s = 0.25 and y/s = 0.40 that 

bounded the right-hand vertical tail. The 

chordwise distributions of the PSP pressure 

coefficient at these stations are presented in the 

plots in the upper left-hand portion of figure 13, 

where Cp is plotted versus the pixel number. 

The first pixel in each row is the closest to the 

wing leading edge, and the last pixel in each row 

is closest to the wing trailing edge. The 

chordwise row at y/s = 0.25 has approximately 

742 pixels, is positioned inboard of the LEX 

vortex pressure footprint and the vertical tail, 

and extends through the high-pressure region 

downstream of the tail shock. The chordwise 

row at y/s = 0.40 has approximately 632 pixels 

and is positioned outboard of the vertical tail, 

intersects the wing leading-edge vortex pressure 

footprint near the leading edge, and extends 

through the suction pressure region induced by 

the hypothesized vortex flow shed from the tail. 

Both rows are identified as black dashed lines in 

the PSP image. The chordwise location of the 

vertical tail apex extended to the two pressure 

rows corresponds to pixel 587 at y/s = 0.25 and 

pixel 477 at y/s = 0.40. Situated directly below 

and scaled with respect to the PSP image is the 

corresponding LVS cross-flow pattern in the 

near wake (x/c = 1.10) of the right-hand wing 

and vertical tail. The PSP pressure distribution 

at y/s = 0.25 is situated near the inboard edge of 

the LEX vortex pressure footprint, and the 

corresponding pressure distribution indicates a 

flow expansion to an approximately constant 

level along the main portion of the wing up to 

the region of the vertical tail. A significant flow 

recompression occurs downstream of the oblique 

shock wave generated at the tail. The estimated 

shock front position corresponds to pixel 599 at 

y/s = 0.25. The rapid pressure rise corresponds 

to the abrupt change in the false-colored PSP 

image from green to yellow to red/pink across 

the shock. The pressure row at y/s = 0.40 

intersects the wing vortex pressure signature 

near the leading edge, which appears as a blue 

region in the PSP pressure map. The pressure 

distribution reveals a corresponding vortexinduced suction pressure peak in this region. 

The pressure row at y/s = 0.40 is situated 

between the wing and LEX vortex pressure 

footprints farther aft, and it captures a region of 

approximately constant pressure level on the 

wing (depicted as green in the pressure map) up 

to the vicinity of the vertical tail. The PSP 

pressure map suggests the presence of a weak 

shock front extending outboard from the tail, 

which is consistent with the minor flow 

recompression shown in the pressure distribution 

at y/s = 0.40. The superposition of a strong 

vortex flow field in this region reduces the shock 

strength, however. Farther aft, the pressure 

distribution reveals a significant increase in the 

suction pressure level and a well-defined suction 

pressure peak that is approximately centrally 

located in a region of blue and purple colors in 

the PSP pressure map. The transition from 

green to blue colors at y/s = 0.40 in the PSP 

image occurs at approximately pixel 485. The 

pressure distribution and false-colored pressure 

map in this region are consistent with the 

induced effects of a vortex flow emanating from 

the vertical tail. The LVS image in the near 

wake of the wing confirms the presence of this 

vortex situated at the outboard junction of the 

tail and wing upper surface. The tail vortex 

appears as a dark region with essentially no 

water vapor condensate surrounded by a lighter 

region of condensation. 

Generic Double Delta Wing 

The PSP and LVS images obtained on the 

double delta wing model at M = 0.85 and 

M = 1.20 are shown in figure 14 corresponding 

to  = 20o

. PSP color bars are not presented for 

the double delta wing since the intensity ratio 

scales and corresponding color coding were 

tailored to each angle of attack and Mach 
11

number to extract important features of the 

surface pressure field response. Specification of 

a common intensity ratio scale for all angles of 

attack and Mach numbers would have resulted in 

large, uninterpretable white or black regions in 

certain images. The comparisons of the PSP 

images are solely intended to illustrate 

significant qualitative differences caused by 

Mach number changes at the selected angles of 

attack. In general, blue, lavender, and purple 

colors correspond to regions of low pressure 

(high suction pressure coefficients), whereas 

regions of higher surface pressures are 

represented by yellow and red colors. A green 

color corresponds to an intermediate suction 

pressure level. The LVS cross-flow patterns are 

presented directly below the PSP images with 

which they are associated. LVS images 

obtained at x/c = 0.25, 0.75, and 0.90 are shown, 

where x is the distance measured along the 

model centerline beginning at the strake apex. 

These locations correspond to pressure 

measurement stations where discrete pressure 

tap data were acquired in spanwise rows along 

the left strake and wing upper surfaces. The 

PSP image at M = 0.85 reveals the pressure 

signature of the strake vortex beginning near the 

strake apex and extending to nearly the 90 

percent chord station. The intensity of the strake 

vortex pressure signature is highest in a region 

beginning near the strake-wing junction to 

approximately the 75 percent chord station. 

This region is denoted by colors ranging from 

lavender to purple. The vortex that is shed from 

the wing leading edge is stronger than the strake 

vortex (reference 12), and it induces significant 

suction pressures over a portion of the wing 

upper surface. However, the vortex feeding 

sheet does not remain attached to the leading 

edge because of the moderate wing sweep and 

the induced effect of the strake vortex. As a 

result, the wing vortex moves inboard and 

upward away from the surface (reference 3), and 

its induced effect on the wing surface pressure 

field diminishes. The strake vortex is no longer 

fed by leading-edge vorticity downstream of the 

strake-wing junction, yet its surface pressure 

footprint is most intense in this region. This 

effect may be caused by a downward 

displacement toward the wing surface as a result 

of its interaction with the wing vortex. 

Downstream of approximately the 90 percent 

chord station, the pressure signatures of the 

strake and wing vortices are diffuse or indistinct. 

This trend is consistent with the onset of vortex 

breakdown upstream of the wing trailing edge. 

The presence of an apparent normal shock wave 

situated between the 75 percent and 90 percent 

chord stations and centered on the wing may 

contribute to the speculated vortex instability. 

The shock presence is inferred from the abrupt 

color change from green to yellow to red, and 

the spanwise extent of the shock may be limited 

by a three-dimensional relief effect induced by 

the presence of strong vortical flows over the 

wing (reference 5). The LVS images at 

x/c = 0.25 and x/c = 0.75 reveal a stable ‘donutshaped’ strake vortex cross section in proximity 

to the upper surface, which correlates with the 

strong pressure signature in the PSP image. The 

discontinuous band of condensate above the 

vortex at x/c = 0.75 is interpreted as a 

discontinuity in the vortex feeding sheet. The 

wing vortex is not visible in the LVS flow 

visualization at M = 0.85 because of insufficient 

local condensation. In contrast, more water 

vapor condensate accumulates within the strake 

vortex making it highly visible. This may be 

related to the long run length along which 

leading-edge vorticity sheds and feeds into the 

strake vortex before it passes over the wing 

upper surface. The onset of vortex breakdown 

inferred from the PSP image is confirmed in the 

LVS cross-flow patterns at x/c = 0.90, where the 

vortex has expanded and the core region has 

filled with water vapor condensate. 

A notable transformation occurs in the PSP 

image at M = 1.20 relative to M = 0.85 as shown 

on the right-hand side of figure 14. Increasing 

the Mach number from 0.85 to 1.20 weakens the 

leading-edge vortices and reduces the adverse 

longitudinal pressure gradient through which the 

vortices must traverse (reference 5). Direct 

vortex interaction and normal shock effects that 

occur at the lower Mach number are mitigated or 

eliminated at the higher Mach number, and the 

vortices will persist in a stable form to higher 

angles of attack. The PSP image at M = 1.20 

captures distinct strake and wing vortex pressure 

signatures that persist to the wing trailing edge. 

Direct interaction, or intertwining, of the strake 
12

and wing vortices does not occur at this Mach 

number. The PSP image suggests the wing 

vortex feeding sheet remains attached to the 

leading edge to nearly the wing tip. Furthermore, 

the normal shock footprint observed at M = 0.85 

is eliminated at M = 1.20. The LVS cross-flow 

patterns at M = 1.20 are consistent with the 

trends in the PSP image and also reveal a 

transformation of the condensation patterns 

compared to M = 0.85. Condensation at the 

higher Mach number first appears in the free 

stream, so the vortical flows are visible as dark 

regions in the absence of scattered light 

surrounded by a light background. The LVS 

cross-flow patterns reveal a juxtaposition of 

stable strake and wing vortices. The strake 

vortex is approximately circular in cross section, 

becoming less so further aft on the wing, 

whereas the wing vortex is a flattened, 

elliptically-shaped region that appears attached 

to the leading edge. The attachment of the 

feeding sheet is assumed when the narrow band 

of condensate along the outer edge of the wing 

vortex is attached to the leading edge. 

Expanded views of the PSP surface 

pressure response on the left-hand strake and 

wing upper surface at M = 0.50 and M = 1.20 are 

presented in figure 15 corresponding to  = 20o

. 

The corresponding LVS cross-flow patterns 

about the left-hand wing at x/c = 0.75 are also 

shown directly below the PSP images. The PSP 

and LVS images are scaled relative to each other 

and, also, to the plots of the spanwise Cp

distributions on the left wing at x/c = 0.75 shown 

at the bottom of figure 15. The pressure 

distributions were obtained using the ESP and 

PSP techniques. Note that the Cp distributions 

are plotted to different scales at M = 0.50 and 

M = 1.20. Figure 15 provides representative 

results obtained at a lower subsonic Mach 

number and at a higher transonic Mach number 

to highlight the different flow topologies 

encountered across the range of Mach number in 

the 8-Foot TPT testing. The PSP image at 

M = 0.50 and  = 20o

 reveals the strake vortex 

pressure signature highlighted in blue-tolavender colors that persists just downstream of 

the 75 percent chord station. The PSP intensity 

field response to the strake vortex is highest near 

the strake-wing intersection, where the strake 

vortex is drawn closer to the wing surface as a 

result of an interaction with the wing vortical 

flow. The wing vortex pressure footprint is 

visible near the strake-wing intersection as a 

region of blue, lavender, and purple colors, but 

this footprint quickly becomes obscured as the 

vortex detaches from the leading edge and 

interacts with the dominant strake vortex. 

Downstream of the 75 percent chord station, the 

intensity of the combined strake and wing vortex 

footprint diminishes, and the expanding region 

of yellow and red colors along the aft portion of 

the wing suggests the onset of vortex 

breakdown. The LVS cross-flow image at 

x/c = 0.75 reveals a stable, donut-shaped strake 

vortex, which is consistent with the PSP 

pressure field response at this station. The LVS 

pattern also shows a cloud of water vapor 

condensate outboard of and above the strake 

vortex, which is interpreted as a burst wing 

vortex flow, thereby obscuring the wing vortex 

footprint in the PSP image. The flow situation 

inferred from the PSP and LVS images at 

M = 0.50 leads to the single-peaked pressure 

distribution at the bottom of figure 15, where the 

spanwise location of the suction pressure peak 

correlates with the positions of the strake vortex 

pressure footprint in the PSP color mapping and 

the condensation pattern in the LVS image. The 

significantly diminished suction pressure level 

along approximately the outer 40 percent of the 

local semispan is attributed to the occurrence of 

wing vortex breakdown. The salient features of 

the PSP image at M = 1.20 and  = 20o

 in 

figure 15 include well-defined strake and wing 

vortex pressure signatures that are independent 

and persist to the wing trailing edge. The 

independence of the signatures implies no direct 

interaction between the vortical flows, and from 

their persistence one can infer stability. The 

character of the pressure signatures suggests the 

wing vortex is very broad compared to the strake 

vortex. Inferences from the PSP image are 

supported by the LVS cross-flow pattern at 

x/c = 0.75, and the flow topology in the PSP and 

LVS images is consistent with the pressure 

coefficient distribution at x/c = 0.75. The 

spanwise Cp distribution shows a pronounced 

suction pressure peak correlating with the 

location of the strake vortex in the PSP and LVS 

images and, farther outboard, a broader 
13

distribution featuring a maximum suction 

pressure plateau that correlates with the flatter, 

elliptically-shaped wing leading-edge vortex. 

Faceted Missile Model 

PSP and LVS images obtained on the 

faceted missile model at M = 1.20 are shown in 

figure 16 corresponding to  = 12o

 and  = 20o. 

The size of the PSP images shown at the top of 

figure 16 are different at the two angles of 

attack, since the facility arc sector support 

mechanism pitched the model closer to the PSP 

image acquisition camera installed in the test 

section ceiling as the angle of attack was 

increased from 12o

 to 20o. The LVS flow 

visualization results correspond to M.S. 20.40, 

25.50, and 30.00 expressed in inches measured 

from the nose tip of the 36-inch-long model. 

Discrete pressure tap data were obtained in 

spanwise rows at these stations using the ESP 

method. The LVS images are scaled relative to 

each other but not to the PSP images. 

The PSP intensity field response is more 

subtle on the faceted missile model compared to 

the 65-degree cropped delta wing-LEX and 

double delta wing models discussed in previous 

sections of this report. The slender forebody, 

forward-swept aftbody, thick diamond-shaped 

cross sections, and large chine included angles at 

the leading edges diminish the strength of the 

vortical flows compared to the thin, sharp-edge, 

less highly-swept wings. In addition, the slender 

missile model is more prone to flow-field 

asymmetries at zero sideslip angle (), which is 

apparent in the PSP images particularly along 

the aftbody region. A PSP color bar is included 

in figure 16 which indicates that regions of 

higher suction pressures are coded green, blue, 

and purple, and regions of lower suction 

pressures or small positive pressures are 

represented by shades of yellow, red, and pink. 

The footprints of the forebody vortices at 

 = 12o are initially discernible in the PSP image 

as narrow green bands along the forebody, 

which intensify to blue color along the transition 

section of the body. The vortex footprints 

appear to widen along the aftbody, albeit in an 

asymmetric fashion relative to the model 

centerline, and the lighter green and yellow 

color coding is indicative of diminished 

influence of the vortices on the surface 

pressures. A small region of dark blue color is 

also apparent near the planform break, which 

suggests a second region of organized flow 

separation. The off-surface flow field at  = 12o

is dominated by the vortex pair shed from the 

forebody as shown in the LVS cross-flow 

patterns in figure 16. The scale of the vortex 

structure in the LVS flow visualization contrasts 

with the relatively subtle surface pressure 

signatures in the PSP image, particularly along 

the aftbody region. A band of condensation 

connects each vortex to the body edge at M.S. 

20.40, which indicates the feeding sheet is 

attached to the leading edge. The body chine is 

essentially a trailing edge downstream of the 

planform break, and it is presumed that vorticity 

is no longer feeding into the vortex from this 

point aft. Consequently, the flow about the 

afterbody section is characterized by a free 

vortex system. An interesting feature of the 

LVS cross-flow patterns at M.S. 25.50 and 

30.00 is the development of a second pair of 

smaller, co-rotating vortices that directly interact 

with the forebody vortical flows. It is speculated 

that a discontinuity exists in the forebody 

leading-edge vortex shear layer upstream of the 

planform break. The premature detachment of 

the forebody vortex feeding sheet from the body 

edge upstream of the planform break allows a 

second pair of co-rotating vortices to form along 

the transition region. A strong interaction of the 

co-rotating vortices occurs along the aftbody as 

shown in the LVS pattern at M.S. 30.00. The 

vortex system is no longer attached to the model 

via a leading-edge feeding sheet in this region 

and, therefore, it migrates away from the upper 

surface. This upward displacement is a 

plausible explanation for the diminished surface 

pressure signature along the aft portion of the 

model. The flow topology in the PSP and LVS 

images at  = 20o

 shown on the right-hand side 

of figure 16 is similar to that observed at 

 = 12o. Increasing the angle of attack amplifies 

the forebody vortex pressure signature and the 

scale of the vortex cross-flow patterns in the 

PSP and LVS images, respectively. The LVS 

condensation patterns are dominated by the 

forebody primary vortices, and the influence of 

the second pair of co-rotating vortices is 
14

significantly diminished at the higher angle of 

attack. The streamwise accumulation of water 

vapor condensate on the long slender model 

creates a vivid depiction of the vortical flows. 

However, the scale and complexity of the 

vortex-dominated condensation patterns along 

the aftbody are incongruous with the relatively 

subtle features in the PSP surface pressure 

response. 

The PSP images and selected LVS crossflow patterns at M = 1.20 corresponding to 

 = 12o and  = 20o in figure 16 are expanded in 

figure 17 for comparison to the upper surface 

static pressure coefficient distributions at M.S. 

25.50. Only the LVS images are appropriately 

scaled to the Cp distributions shown at the 

bottom of the figure. Straightforward 

correlations are made between the spanwise 

locations of the forebody primary vortices and 

the suction pressure peaks at  = 12o

 and 

 = 20o. The presence of the second pair of corotating vortices at  = 12o

 is not manifested in 

the pressure distributions, although they are 

clearly visible in the LVS image. The influence 

of the relatively large regions of secondary 

separation denoted in the LVS image at  = 20o

is associated with inflection points in the 

pressure distributions outboard of the primary 

vortex suction peaks. 

Figure 18 shows the effect of sideslip on 

the PSP surface pressure response and LVS 

cross-flow patterns at M = 0.90 and  = 20o

. 

The forebody primary vortex pressure signatures 

at  = 0o

 appear as narrow bands of blue color 

that are symmetrically positioned along the 

forebody and broader bands of green color that 

are asymmetrically disposed along the rear 

portion of the aftbody. A secondary region of 

locally high suction pressures is coded dark blue 

in the transition section of the body, which is 

similar to the pattern observed at M = 1.20 and 

 = 20o in figure 16 and figure 17. The LVS 

patterns at  = 0o

 reveal a symmetric pair of 

forebody vortices at M.S. 20.40 and the 

development of a second pair of smaller vortices 

rotating in the same sense at M.S. 25.50 that is 

presumed to originate from the transition region. 

The latter vortex pair is entrained into the 

dominant forebody vortices, and they appear as 

spiral discontinuities at the outer edge of the 

vortex cross-flow images at M.S. 30.00. As a 

consequence, the presence of the co-rotating 

vortex pair is not manifested in the PSP image. 

There is some evidence of vortex asymmetry at 

M.S. 25.50 and 30.00, where the right-hand 

vortex system is situated slightly higher above 

the surface. The PSP image at  = -2.5o

 (nose 

right) in figure 18 reveals a pressure signature of 

the left-hand, or windward, forebody vortex, 

which is discernible as a narrow band of blue 

and dark green colors over most of the body 

length. In addition, there is a second region of 

higher suction pressures coded dark blue/purple 

on the windward side along the transition region 

of the body. In contrast, the pressure signature 

of the forebody vortex on the right-hand, or 

leeward, side is less visible along the forebody 

region and barely discernible along the aftbody. 

Good qualitative correlation exists between the 

PSP and LVS images. At M.S. 20.40, for 

example, the windward forebody vortex is close 

to the surface and is attached to the leading edge 

by a band of condensate. The leeward vortex is 

situated farther above the surface, and there is no 

evidence of a feeding sheet connecting the 

vortex to the leading edge at this model station. 

In sideslip, the windward vortex moves closer to 

the model surface while the leeward vortex 

migrates away from the surface (reference 4). 

This effect contributes to the higher PSP suction 

pressure levels on the windward side compared 

to the leeward side. In addition, the effective 

decrease in the leading-edge sweep on the 

windward side yields a stronger vortex at a 

given angle of attack, whereas the opposite 

effect occurs on the leeward side (reference 12). 

The development of a second, co-rotating vortex 

is apparent at M.S. 25.50 from the transition 

region of the body on the windward side. This 

smaller, weaker vortex is observed to orbit about 

the dominant forebody vortex at M.S. 30.00, 

which remains in proximity to the model 

surface. The latter trend is consistent with the 

persistence of the vortex pressure signature in 

the PSP image. By comparison, the apparent 

free vortex system on the leeward side induces a 

more subtle pressure signature. 
15

The PSP false-colored pressure maps and 

the LVS images at M.S. 25.50 from figure 18 

are expanded in figure 19 for comparison to the 

spanwise Cp distributions at M.S. 25.50 

corresponding to  = 0o

 and  = -2.5o. The LVS 

images are approximately scaled to their 

respective Cp distribution plots. The locations 

of the bands of green/blue color in the PSP 

pressure map and the forebody vortex cross-flow 

patterns in the LVS image at  = 0o

 correlate 

with the suction pressure peak locations in the 

Cp plot at the bottom left of figure 19. The 

aftbody flow asymmetry noted in figure 18 at 

 = 0o is manifested as a slight asymmetry in the 

magnitude of the peak Cp values. The presence 

of the second pair of co-rotating vortices visible 

in the LVS cross-flow pattern does not project to 

the Cp plot, because of their spatial locations 

relative to the model surface and the dominance 

of the forebody vortices. The significant 

asymmetries that are imposed on the PSP and 

LVS images due to sideslip correlate very well 

with the corresponding Cp plot at  = -2.5o

 at the 

bottom right of figure 19. The stronger 

windward vortex in proximity to the upper 

surface induces a much higher suction pressure 

peak compared to its weaker leeward 

counterpart that is positioned higher above the 

model. 

Concluding Remarks 

Laser vapor screen flow visualization and 

pressure-sensitive paint techniques were applied 

in a tandem approach to investigate the vortexand shock wave-dominated flow fields about 

slender wing and missile configurations at 

subsonic, transonic, and supersonic speeds. 

Selected results were presented from 

experiments performed on 65-degree cropped 

delta wing-LEX, double delta, and faceted 

missile models in the NASA Langley Research 

Center Unitary Plan Wind Tunnel and the 8-Foot 

Transonic Pressure Tunnel. The LVS technique 

was considered qualitative, since the appearance 

of condensation in the test section affects the 

free-stream flow characteristics. The potential 

intrusive effects of PSP were mitigated by 

controlling the thickness and surface roughness 

of the luminescent paint coating. Within these 

constraints, however, the LVS cross-flow 

patterns were effectively projected to the falsecolored PSP surface pressure maps to identify 

the salient features of the aerodynamic flows 

about the selected models. The LVS and PSP 

techniques were complementary in that they 

provided a more complete depiction of the 

complex three-dimensional flow fields that were 

characterized by the development and 

interaction of multiple leading-edge vortices and 

shock waves, vortex breakdown, and vortex 

interactions with vertical tail surfaces. The 

combination of these techniques was also a 

powerful tool in visualizing and quantifying the 

effects of passive surface porosity on vortex 

flow development and interactions. The 

application of the LVS and PSP techniques was 

mutually exclusive, because of the different 

requirements for the optical components, model 

surface preparation, and wind tunnel test 

conditions. Consequently, the LVS and PSP 

results were obtained in separate phases in each 

experiment, which required appropriate test 

planning and efficient use of wind tunnel facility 

resources. The successive application of these 

techniques often yielded synergistic results, 

however, since certain flow features that were 

subtle or not discernible using one technique 

were often clearly revealed using the other. 

References 

1. Erickson G E. Overview of Supersonic 

Aerodynamics Measurement Techniques in 

the NASA Langley Unitary Plan Wind 

Tunnel. NASA TM-2007-214894, 2007.

2. Erickson G E. Wind Tunnel Investigation of 

the Effects of Surface Porosity and Vertical 

Tail Placement on Slender Wing Vortex 

Flow Aerodynamics at Supersonic Speeds. 

NASA TM-2007-215082, 2007. 

3. Erickson G E and Gonzalez H A. Wind 

Tunnel Application of a Pressure-Sensitive 

Paint Technique to a Double Delta Wing 

Model at Subsonic and Transonic Speeds. 

NASA TM-2006-214319, 2006. 

4. Erickson G E. Wind Tunnel Application of a 

Pressure-Sensitive Paint Technique to a 

Faceted Missile Model at Subsonic and 
16

Transonic Speeds. NASA TM-2004-

212991, 2004. 

5. Erickson G E. Wind Tunnel Investigation of 

the Interaction and Breakdown 

Characteristics of Slender-Wing Vortices at 

Subsonic, Transonic, and Supersonic 

Speeds. NASA TP-3114, 1991. 

6. Amer T R, Liu T and Oglesby D M. 

Characterization of Pressure Sensitive Paint 

Intrusiveness Effects on Aerodynamic Data. 

AIAA 2001-0556, 39th AIAA Aerospace 

Sciences Meeting & Exhibit, Reno, NV, 

2001. 

7. Jackson C M, Corlett W A and Monta W J. 

Description and Calibration of the Langley 

Unitary Plan Wind Tunnel. NASA TP 

1905, 1981. 

8. Brooks C W Jr., Harris C D and Reagon P 

G. The NASA Langley 8-Foot Transonic 

Tunnel Calibration. NASA TP 3437, 1994. 

9. Elsenaar A. How It All Started: The 

International Vortex Flow Experiment on 

Euler Code Validation in Retrospect. 

Symposium on International Vortex Flow 

Experiment on Euler Code Validation – 

Proceedings, A. Elsenaar and G. Eriksson, 

eds., FFA, Flygtekniska Forsoksanstalten 

(Sweden), Oct. 1986, pp. 17-19. 

10. Kern S B. Numerical Investigation of 

Vortex Flow Control Through Small 

Geometry Modifications at the Strake/Wing 

Junctions of a Cropped Double Delta Wing. 

AIAA-92-0411, January 1992. 

11. Wood R M. Supersonic Aerodynamics of 

Delta Wings. NASA TP 2771, 1988. 

12. Hemsch M J and Luckring J M. Connection 

Between Leading-Edge Vortex Lift and 

Vortex Strength for Delta Wings. Journal of 

Aircraft, Vol. 27, No. 5, pp. 473-375, 1990. 
17 

Figure 1. Laser vapor screen flow visualization at M = 1.6, = 20 degrees (reference 5). 

 Photograph was obtained from a one-quarter left front view; laser light sheet 

 originated from opposite side of wind tunnel test section. 

Figure 2. Pressure-sensitive paint surface pressure response on a slender wing at M = 1.65 and 

= 6 degrees (reference 1). 

vortex flow 

shock wave 

shock wave 

vortex flow

vortex pressure 

signature 
18

Figure 3. Photographs of the 65-degree cropped delta wing-LEX model installed in Test 

 Section 1 of the NASA Langley Research Center Unitary Plan Wind Tunnel 

 (photograph to the right shows the model prepared for PSP testing). 

(a) planview 

Figure 4. Details of the 65-degree cropped delta wing-LEX model (all dimensions in inches). 

 (B.L. = butt line, BMC = balance moment center, M.S. = model station, 

 MRC = moment reference center, W.L. = water line) 

Moment

Reference

Center (MRC) Wing-Mounted

Vertical Tail

Centerline

Vertical Tail

A

A

Section A-A

ESP Module

M.S.

0.00

M.S.

7.68

M.S.

14.767

M.S.

21.853

M.S. 

26.578

M.S.

27.758

M.S.

31.302

M.S.

32.286

B.L.

9.360

B.L.

3.120

B.L.

-1.575

M.S.

21.144

B.L.

2.480

Porous

LEX

B

B

C

C

65°

65°

Section B-B

45°

0.25

Section C-C

LEX

Beveled Wing

leading edge

90°

90°
19

(b) sideview with centerline vertical tail 

Figure 4. Continued. 

(c) sideview with twin vertical tails 

Figure 4. Concluded. 

M.S.

0.00

M.S. 

21.144

M.S. 

31.302

M.S. 

32.286

UT-52A BMC

M.S. 

26.832

M.S. 

31.602

M.S. 

32.482

W.L. 

5.714

W.L. 

0.984

W.L. 

0.000

MRC

45.24°

M.S. 

29.598

W.L. 

2.773

10 psid 48-port 

ESP purge module

M.S.

0.00

M.S. 

21.144

M.S. 

31.302

M.S. 

32.286

350-5B Sting

Gooseneck

LEX

M.S. 

25.921

M.S. 

31.750

M.S. 

34.924

W.L. 

7.404

W.L. 

1.575

W.L. 

0.984

W.L. 

0.000

M.S. 

32.802

 1.375

 2.500

3.150

70

Rear view of

model base

45°

MRC

M.S. 

29.791

M.S.

21.475 UT-52A BMC

W.L. 

4.131
20

Figure 5. Photographs of the double delta wing model installed in the test section of 

 the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel 

 (photograph to the right shows the model prepared for PSP testing). 

Figure 6. Planview and sideview of the double delta wing model (all dimensions in inches). 

76°

12.000

4.000

2.084

13.460

8.359

11.723

16.326

MRC Balance

Sting

Pressure Row

ESP Module

A

A

B B

14.400

40°

16.000

Baseline Fillet

Sections A-A and B-B

R.005 Continuous

0.375 20°

1.030

16.000

11.723

4.400

17.600

0.375

2.200

1.600

A

A

Section A-A

R1.000

Fiberglass housing
21

Figure 7. Photographs of the faceted missile model installed in the test sections of NASA 

 Langley Research Center wind tunnels (photograph to the right shows the 

 model in the NASA LaRC 8-Foot TPT prepared for PSP testing). 

Figure 8. Details of the faceted missile model (all dimensions in inches). 

MRC 
22

Figure 9. Comparison of PSP and LVS images on the 65-degree cropped delta wing model 

 with solid LEX and porous LEX at M = 1.6,  = 8o

 (from NASA LaRC UPWT). 

Solid LEX

M=1.6 α=8o Porous LEX

M=1.6 α=8o

x/c=0.80

x/c=0.60

x/c=1.00

LEX vortex 2 wing vortex

1

2 2

1 3 tail shock 5 tip vortex

3

2

2

5 3 5

2

3

1

2

x/c=0.80

x/c=0.60

x/c=1.00

+.10 −.33

Cp level

3

1

4 4

4 LEX−wing junction
23

Figure 10. Correlation of PSP and LVS images with spanwise pressure distributions on the 65-degree cropped 

 delta wing model with solid LEX and porous LEX at M = 1.6,  = 8o

 (from NASA LaRC UPWT). 

MACH=1.60, AOA=8 DEGREES, X/C=0.80

0.0 0.2 0.4 0.6 0.8 1.0

Y/S

0.05

0.00

-0.05

-0.10

-0.15

-0.20

-0.25

-0.30

-0.35

CP

METHOD

ESP

PSP

MACH=1.60, AOA=8 DEGREES, X/C=0.80

0.0 0.2 0.4 0.6 0.8 1.0

Y/S

0.05

0.00

-0.05

-0.10

-0.15

-0.20

-0.25

-0.30

-0.35

CP

METHOD

ESP

PSP

LEX vortex

wing vortex

Solid LEX

M=1.6 α=8o

x/c=0.80

Porous LEX

M=1.6 α=8o

1

2 2

1

2 2

2

1

+.10 −.33

Cp level

3 3

LEX−wing

junction

3

x/c=0.80

x/c=0.80
24

Figure 11. Comparison of LVS images on the 65-degree cropped delta wing model with 

 solid LEX and porous LEX at M = 2.1,  = 16o

 (from NASA LaRC UPWT). 

Solid LEX

M=2.1 α=16o

x/c = 0.80

1

2

3

4

1 LEX vortex 2 wing vortex 3 cross−flow shock

4 Combined LEX and wing vortex

Porous LEX 3

M=2.1 α=16o

x/c = 0.80

3
25

Figure 12. Comparison of PSP and LVS images on the 65-degree cropped delta wing model with solid 

 LEX and centerline and twin vertical tails at M = 1.6,  = 8o

 (from NASA LaRC UPWT). 

Solid LEX

CenterlineTail

M=1.6 α=8o

Solid LEX

Twin Tails

M=1.6 α=8o

x/c=0.80

x/c=0.60

x/c=1.10

1 2 1

2

1

2

1

2

4

6

5 2

4

1 LEX vortex 2 wing vortex 3 tail shock 4 tip vortex Combined LEX and wing vortex

6 wake rollup tail vortex

5

7

7

3 3

7

+.10 −.33

Cp level

8 8

8 LEX−wing junction

3 1

x/c=1.10

x/c=0.80

x/c=0.60

3
26

Figure 13. Correlation of PSP and LVS images with chordwise pressure distributions on the 65-degree cropped delta 

 wing model with solid LEX and twin vertical tails at M = 1.6,  = 8o

 (from NASA LaRC UPWT). 

y/s=0.25

y/s=0.40

Mach=1.60, y/s=0.40, Solid LEX

0 100 200 300 400 500 600 700

Pixels

0.3

0.2

0.1

0.0

-0.1

-0.2

-0.3

-0.4

-0.5

Cp

Mach=1.60, y/s=0.25, Solid LEX

0 100 200 300 400 500 600 700 800

Pixels

0.3

0.2

0.1

0.0

-0.1

-0.2

-0.3

-0.4

-0.5

Cp

y/s=0.25

y/s=0.40

1

2

3

2

5

1 4

LEX vortex wing vortex

3 tail shock tip vortex

tail vortex

+.10 −.33

Cp level

7

5

6

weak shock front

7 LEX−wing junction

2

5 6

1

4

x/c=1.10

3
27

Figure 14. Comparison of PSP and LVS images on the double delta wing model at 

 M = 0.85 and M = 1.20,  = 20o

 (from NASA LaRC 8-Foot TPT). 

x/c=0.25

x/c=0.75

x/c=0.90

x/c=0.25

x/c=0.75

x/c=0.90

1 strake vortex 2 wing vortex 3 normal shock

M=0.85

α=20o

M=1.20

α=20o

1 1

2 2

4 4

4 vortex feeding

sheet discontinuity

5 vortex breakdown

3

1 4 1 2

5 1 2

1 1
28

Figure 15. Correlation of PSP and LVS images with spanwise pressure distributions on the double delta 

 wing model at M = 0.50 and M = 1.20,  = 20o

 (from NASA LaRC 8-Foot TPT). 

MACH=0.50, AOA=20 DEGREES, X/C=0.75

-1.0 -0.8 -0.6 -0.4 -0.2 0.0

Y/S

0.4

0.0

-0.4

-0.8

-1.2

-1.6

-2.0

CP

METHOD

ESP

PSP

MACH=1.20, AOA=20 DEGREES, X/C=0.75

-1.0 -0.8 -0.6 -0.4 -0.2 0.0

Y/S

0.0

-0.2

-0.4

-0.6

-0.8

-1.0

CP

METHOD

ESP

PSP

strake

vortex

wing

vortex

M=0.50

α=20o

M=1.20

α=20o

2

1

wing vortex

breakdown

3 1 2

1

2

2

3

1

x/c=0.75

x/c=0.75

x/c=0.75

1
29

Figure 16. Comparison of PSP and LVS images on the faceted missile model at 

 = 12o and  = 20o, M = 1.20 (from NASA LaRC 8-Foot TPT). 

M.S. 20.40

M.S. 25.50

M.S. 30.00

M.S. 20.40

M=1.20

α=12o M=1.20

α=20o

M.S. 25.50

M.S. 30.00

1 forebody vortex 2 secondary separation 3 co−rotating vortex

1

3 3

1

1

3

1

3

1

3

3

1

1

2

1

2
30

Figure 17. Correlation of PSP and LVS images with spanwise pressure distributions on the faceted 

 missile model at  = 12o

 and  = 20o, M = 1.20 (from NASA LaRC 8-Foot TPT). 

M=1.20

α=12o

M=1.20

α=20o

1 forebody vortex 2 secondary separation 3 co−rotating vortex

Mach=1.20, AoA=12 degrees, M.S. 25.50

-1.00 -0.75 -0.50 -0.25 0.00 0.25 0.50 0.75 1.00

y/s

0.0

-0.1

-0.2

-0.3

-0.4

-0.5

-0.6

Cp

ESP

PSP

Mach=1.20, AoA=20 degrees, M.S. 25.50

-1.00 -0.75 -0.50 -0.25 0.00 0.25 0.50 0.75 1.00

y/s

0.0

-0.1

-0.2

-0.3

-0.4

-0.5

-0.6

Cp

ESP

PSP

1 1

3 3

1

2

3

3

1

M.S. 25.50

M.S. 25.50

M.S. 25.50
31

Figure 18. Comparison of PSP and LVS images on the faceted missile model at 

 = 0o and  = -2.5o, M = 0.90,  = 20o (from NASA LaRC 8-Foot TPT).

M.S. 20.40

M.S. 25.50

M.S. 30.00

M.S. 20.40

M=0.90

α=20o β=0o

M=0.90

α=20o β=−2.5o

M.S. 25.50

M.S. 30.00

1

5

2

1 forebody vortex 2 secondary separation 3 co−rotating vortex

4 windward forebody vortex 5 leeward forebody vortex

4

1

1

3

4 5

3

3 3

1
32

Figure 19. Correlation of PSP and LVS images with spanwise pressure distributions on the faceted 

 missile model at  = 0o

 and  = -2.5o, M = 0.90,  = 20o (from NASA LaRC 8-Foot TPT). 

M=0.90

α=20oβ=0o M=0.90

α=20o β=−2.5o

M.S. 25.50

5

1 forebody vortex 2 secondary separation 3 co−rotating vortex

windward forebody vortex leeward forebody vortex

4

Mach=0.90, AoA=20 degrees, M.S. 25.50

-1.00 -0.75 -0.50 -0.25 0.00 0.25 0.50 0.75 1.00

y/s

0.0

-0.1

-0.2

-0.3

-0.4

-0.5

-0.6

-0.7

-0.8

-0.9

Cp

ESP

PSP

M=0.90, AoA=20 deg, Beta=-2.5 deg, M.S. 25.50

-1.00 -0.75 -0.50 -0.25 0.00 0.25 0.50 0.75 1.00

y/s

0.0

-0.1

-0.2

-0.3

-0.4

-0.5

-0.6

-0.7

-0.8

-0.9

Cp

ESP

PSP

1

3 3

5 1 4

3

2 3

4 5
REPORT DOCUMENTATION PAGE Form Approved

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Technical Memorandum

 4. TITLE AND SUBTITLE

Unified Application of Vapor Screen Flow Visualization and Pressure 

Sensitive Paint Measurement Techniques to Vortex- and Shock 

Wave-Dominated Flow Fields 

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Erickson, Gary E.

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

NASA Langley Research Center

Hampton, VA 23681-2199

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National Aeronautics and Space Administration

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19a. NAME OF RESPONSIBLE PERSON

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14. ABSTRACT

Laser vapor screen (LVS) flow visualization and pressure sensitive paint (PSP) techniques were applied in a unified approach 

to wind tunnel testing of slender wing and missile configurations dominated by vortex flows and shock waves at subsonic, 

transonic, and supersonic speeds. The off-surface cross-flow patterns using the LVS technique were combined with global 

PSP surface static pressure mappings to characterize the leading-edge vortices and shock waves that coexist and interact at 

high angles of attack. The synthesis of LVS and PSP techniques was also effective in identifying the significant effects of 

passive surface porosity and the presence of vertical tail surfaces on the flow topologies. An overview is given of LVS and 

PSP applications in selected experiments on small-scale models of generic slender wing and missile configurations in the 

NASA Langley Research Center (NASA LaRC) Unitary Plan Wind Tunnel (UPWT) and 8-Foot Transonic Pressure Tunnel 

(8-Foot TPT). 

15. SUBJECT TERMS

Aerodynamics; Flow visualization; Laser vapor screen; Pressure sensitive paint; Shock waves; Subsonic; Supersonic; 

Transonic; Vortex flows; Wind tunnel tests

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