April 2013

NASA/TM–2013-217982

Wind Tunnel Investigation of Passive Vortex 

Control and Vortex-Tail Interactions on a 

Slender Wing at Subsonic and Transonic Speeds

Gary E. Erickson

Langley Research Center, Hampton, Virginia
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April 2013

NASA/TM–2013-217982

Wind Tunnel Investigation of Passive Vortex 

Control and Vortex-Tail Interactions on a 

Slender Wing at Subsonic and Transonic Speeds

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 author would like to thank the research and technical staff assigned to the 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

Model Description and Test Apparatus ....................................................................................................4

Wind Tunnel Facilities and Test Conditions.............................................................................................7

Experimental Techniques...........................................................................................................................8

Laser Vapor Screen Technique.................................................................................................................8

ESP Measurements .................................................................................................................................10

Strain Gage Balance Measurements .......................................................................................................10

Model Chamber Pressure Measurements................................................................................................11

Pitch Angle Measurements.....................................................................................................................11

Yaw Angle Measurements......................................................................................................................12

Corrections..............................................................................................................................................12

Flow angularity ...................................................................................................................................12

Other corrections.................................................................................................................................12

Chamber pressures..............................................................................................................................12

Discussion of Results...............................................................................................................................12

Reynolds Number Assessment................................................................................................................13

Pressure Distributions.........................................................................................................................13

Force and Moment Measurements......................................................................................................14

Within-Test Data Repeatability ..............................................................................................................15

Pressure Distributions.........................................................................................................................15

Force and Moment Measurements......................................................................................................16

Tunnel-to-Tunnel Data Reproducibility..................................................................................................17

Pressure Distributions.........................................................................................................................17

Force and Moment Measurements......................................................................................................18

Basic Configuration Aerodynamics........................................................................................................20

Pressure Distributions.........................................................................................................................20

Wing Alone.....................................................................................................................................20

Mach Number Sweeps................................................................................................................20

Angle-of-Attack Sweeps.............................................................................................................21

Sideslip Sweeps ..........................................................................................................................21

Solid LEX .......................................................................................................................................22

Mach Number Sweeps................................................................................................................22

Angle-of-Attack Sweeps.............................................................................................................22

Sideslip Sweeps ..........................................................................................................................23

Porous LEX.....................................................................................................................................23

Mach Number Sweeps................................................................................................................23

Angle-of-Attack Sweeps.............................................................................................................23

Sideslip Sweeps ..........................................................................................................................24

Canard.............................................................................................................................................24

Mach Number Sweeps................................................................................................................24

Angle-of-Attack Sweeps.............................................................................................................25

Sideslip Sweeps ..........................................................................................................................25

Force and Moment Measurements......................................................................................................25

Wing Alone.....................................................................................................................................25

Mach Number Effects.................................................................................................................25
vi

Sideslip Effects ...........................................................................................................................26

Solid LEX .......................................................................................................................................26

Mach Number Effects.................................................................................................................26

Sideslip Effects ...........................................................................................................................27

Porous LEX.....................................................................................................................................27

Mach Number Effects.................................................................................................................27

Sideslip Effects ...........................................................................................................................27

Canard.............................................................................................................................................27

Mach Number Effects.................................................................................................................27

Sideslip Effects ...........................................................................................................................28

Effect of Full LEX Porosity....................................................................................................................28

Laser Vapor Screen Flow Visualization .............................................................................................28

Mach = 0.85 ....................................................................................................................................28

Pressure Distributions.........................................................................................................................29

Mach = 0.50 ....................................................................................................................................29

Mach = 0.85 ....................................................................................................................................30

Mach = 1.20 ....................................................................................................................................30

Force and Moment Measurements......................................................................................................31

Angle-of-Attack Sweeps at Mach = 0.50, 0.85, and 1.20 and β = 0 Degrees.................................31

Mach = 0.50 ................................................................................................................................31

Mach = 0.85 ................................................................................................................................31

Mach = 1.20 ................................................................................................................................32

Sideslip Sweeps at Mach = 0.50, 0.85, and 1.20 and Selected Angles of Attack ...........................32

Mach = 0.50 ................................................................................................................................32

Mach = 0.85 ................................................................................................................................32

Mach = 1.20 ................................................................................................................................33

Full Versus Partial LEX Porosity............................................................................................................33

Pressure Distributions.........................................................................................................................33

Porosity Levels 1, 2 and 3...............................................................................................................33

Porosity Levels 1, 4 and 5...............................................................................................................33

Force and Moment Measurements......................................................................................................34

Porosity Levels 1, 2 and 3...............................................................................................................34

Porosity Levels 1, 4 and 5...............................................................................................................34

Comparison of Vertical Tail Arrangements............................................................................................34

Pressure Distributions.........................................................................................................................34

Mach = 0.50, β = 0 Degrees............................................................................................................34

Mach = 0.85, β = 0 Degrees............................................................................................................35

Mach = 0.50, β = +/-4 and +/-8 Degrees.........................................................................................36

Mach = 0.85, β = +/-4 and +/-8 Degrees.........................................................................................36

Force and Moment Measurements......................................................................................................36

Angle-of-Attack Sweeps at Mach = 0.50 and 0.85 and β = 0 Degrees...........................................36

Mach = 0.50 ................................................................................................................................36

Mach = 0.85 ................................................................................................................................37

Sideslip Sweeps at Mach = 0.50 and 0.85 and Selected Angles of Attack .....................................37

Mach = 0.50 ................................................................................................................................37

Mach = 0.85 ................................................................................................................................37

Comparison of Solid LEX and Canard ...................................................................................................38

Pressure Distributions.........................................................................................................................38

Mach = 0.50 ....................................................................................................................................38
vii

Mach = 0.85 ....................................................................................................................................39

Force and Moment Measurements......................................................................................................39

Angle-of-Attack Sweeps at Mach = 0.50 and 0.85 and β = 0 Degrees...........................................39

Mach = 0.50 ................................................................................................................................39

Mach = 0.85 ................................................................................................................................39

Sideslip Sweeps at Mach = 0.50 and 0.85 and Selected Angles of Attack .....................................40

Mach = 0.50 ................................................................................................................................40

Mach = 0.85 ................................................................................................................................40

Concluding Remarks...............................................................................................................................40

References................................................................................................................................................41

Tables .............................................................................................................................................42

Figures .............................................................................................................................................46
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viii
1

Abstract

A wind tunnel experiment was conducted in the NASA Langley Research 

Center 8-Foot Transonic Pressure Tunnel to determine the effects of 

passive surface porosity on vortex flow interactions about a general 

research fighter configuration at subsonic and transonic speeds. Flowthrough porosity was applied to a wing leading-edge extension, or LEX, 

mounted to a 65-degree cropped delta wing model as a longitudinal 

instability mitigation technique at high angles of attack. In addition, the 

porosity of the LEX was compartmentalized to determine the sensitivity 

of the vortex-dominated aerodynamics to the location and level of 

porosity applied to the LEX. Test data were obtained with LEX on and 

off in the presence of a centerline vertical tail and twin, wing-mounted 

vertical fins having 0-degree and 30-degree cant angles to quantify the 

sensitivity of the vortex flow aerodynamics to tail placement and 

orientation. A close-coupled canard was also tested as an alternative to 

the wing LEX as a high angle-of-attack passive flow control device. 

Wing upper surface static pressure distributions and six-component 

forces and moments were obtained at free-stream Mach numbers of 0.50, 

0.85, and 1.20, unit Reynolds number of 2.5(106

) per foot, angles of 

attack up to approximately 30 degrees, and angles of sideslip to +/-8 

degrees. The off-surface flow field was visualized in cross planes on 

selected configurations using a laser vapor screen flow visualization 

technique. A subset of data was obtained on selected configurations at 

Mach = 0.50 to identify Reynolds number effects on the surface pressure 

distributions and overall force and moment characteristics. An analysis 

of within-test data repeatability was performed. Detailed comparisons 

were also made of the pressure distributions and aerodynamic force and 

moment coefficients obtained on several configurations tested at 

Mach = 0.50 in the 8-Foot Transonic Pressure Tunnel and in a previous 

entry using the same model in the NASA Langley Research Center 

7- by 10-Foot High Speed Tunnel.

Introduction

The control of vortex flows is an important 

design consideration for military and 

commercial aircraft, missiles, and reusable 

launch vehicles. Wings, bodies, stabilizing and 

control surfaces, engine inlets, and surface 

protuberances are several sources where 

controlled flow separation in the form of 

vortices can occur. Disorganized and/or 

unsteady flow separation issues are often 

addressed by adding devices ranging from small 

vortex generators to energize the local boundary 

layer to larger wing leading-edge extensions, 

strakes, and canards to control the global flow 

field. The aerodynamic benefits of vortex flows 

can be extended, and undesired effects 

mitigated, by tailoring the vehicle geometry or 

incorporating passive and active flow control 

concepts. The vortex flow topology can be 

significantly affected by the placement of a 

vertical tail(s) on the fuselage or wings if the 

vortices are in proximity to, or directly interact 

with, the tail surfaces. Passive porosity has been 

successfully applied to control vortices shed 

from slender bodies at subsonic through 

supersonic speeds (reference 1) and to mitigate 

the adverse effects of shock waves on wings at 

transonic and supersonic speeds (reference 2). 

The present experiment focuses on flow-through 

porosity applied to the leading-edge extension

(LEX) of a 65-degree cropped delta wing to 

control the leading-edge vortex development and 

interactions at subsonic and transonic speeds.
2

Partitioning of the porous regions on the LEX 

was also investigated. Figure 1 shows a 

photograph of several of the key wind tunnel 

model components that were used in the current 

experiment. This model was selected since it 

was representative of fighter aircraft designs 

with subsonic/transonic maneuver and 

supersonic cruise capabilities. The model was 

modular in that it could be tested with a LEX, a 

close-coupled canard, and in an isolated wing

configuration. In addition, it had provisions to 

mount a centerline tail and twin, wing-mounted 

uncanted and canted vertical tails, which were of 

interest to this class of fighter aircraft. Porosity 

was applied to the LEX, since it was situated 

ahead of the moment reference center (MRC) 

and generated a strong vortex flow that affected 

the global wing flow field. Primary model

configurations that were tested included the nonporous (solid) LEX, the porous LEX with five 

levels of flow-through porosity, a close-coupled 

canard, and the isolated wing. All of these 

configurations were tested in combination with a 

centerline vertical tail. The solid LEX was also 

tested with twin wing-mounted uncanted and 

canted vertical fins and with no tail. This test 

matrix provided a sufficient range of 

configurations to assess the effects of LEX 

porosity on the wing upper surface static 

pressure distributions and the aerodynamic force 

and moment characteristics; to assess the 

sensitivity of the vortex-dominated 

aerodynamics to the vertical tail placement and 

orientation; and to compare the high angle-ofattack aerodynamics of LEX and canard 

configurations at subsonic and transonic speeds. 

The primary focus of this report is the

interpretation of the wing upper surface static 

pressure distributions and the overall sixcomponent force and moment aerodynamic

coefficients as a function of the various 

configuration changes, the angles of attack and 

sideslip, and the Mach number. A laser vapor 

screen (LVS) flow visualization method was 

used to assist in the identification of flow 

mechanisms and the aerodynamic effects of 

LEX porosity at transonic speeds. Secondary 

objectives of this experiment included an 

assessment of Reynolds number effects, withintest data repeatability, and facility-to-facility 

data reproducibility. The wind tunnel testing 

was performed in the NASA Langley Research 

Center (NASA LaRC) 8-Foot Transonic 

Pressure Tunnel (8-Foot TPT) in Test 1057 at 

free-stream Mach numbers of 0.50, 0.85, and 

1.20 at a unit Reynolds number of 2.5 million 

per foot. A small subset of data was acquired on 

selected configurations at Mach = 0.50 and unit 

Reynolds numbers of 1.5, 2.5, and 3.5 million 

per foot. Within-test data repeatability was 

assessed by comparing data scatter in repeat runs 

to the pressure transducer and six-component 

balance instrumentation calibration accuracies. 

Tunnel-to-tunnel data reproducibility was 

examined by comparing test results obtained at 

Mach = 0.50 on several configurations using the

same model in the NASA LaRC 8-Foot TPT and 

the NASA LaRC 7- by 10-Foot High Speed 

Tunnel (7- by 10-Foot HST).

Nomenclature

ARPBA1.1 chamber area used to compute 

axial force correction due to 

average chamber pressure 

measurement, 4.9090 square 

inches (sq. in.)

AVI audio video interleave

b span distance, inches (in.)

wb reference wing span, 18.726 in.

(also BSPAN1)

B. L. model butt line, in.

BMC balance moment center, 

M.S. 21.811

c wing centerline chord, 

23.622 in. (also CHORD1)

c mean aerodynamic chord, in.

wc wing mean aerodynamic chord, 

16.056 in.

CA axial force coefficient, 

( w )

Axial Force

q S ∞

CD drag force coefficient, 

( w )

Drag

q S ∞
3

Cl rolling moment coefficient, 

 

( w w )

Rolling Moment

qSb ∞

CL lift coefficient, ( w )

Lift Force

q S ∞

Cm pitching moment coefficient, ( w )

Pitching Moment

qSc ∞

Cn yawing moment coefficient, 

 

( w w )

Yawing Moment

qSb ∞

CN normal force coefficient, ( w )

Normal Force

q S ∞

CY side force coefficient, ( w )

Side Force

q S ∞

Cp static pressure coefficient, ( pp q − ∞ ∞ ) Cp,u upper surface static pressure 

coefficient

Cp,v vacuum pressure coefficient, 2 − 2 M ∞ γ * Cp pressure coefficient

corresponding to the local speed 

of sound,(

)







−





+

−

∞

+

∞

1

1

2

1

2

3.5

2

2

γ

γ

γ

M

M

DESL data engineering scripting 

language

ESP electronically

-scanned pressure

0.25c

l tail length, distance between 

25% tail mean aerodynamic 

chord to MRC, in.

LEX leading-edge extension

LVS laser vapor screen M∞ free-stream Mach number (also 

Mach or M)

MRC moment reference center, 

M.S. 21.144

M.S. model station, in.

p local static pressure, pounds per 

square foot (psf)

pcham1 chamber pressure #1, psf

pcham2 chamber pressure #2, psf

p

0 stagnation pressure, psf

p

∞ free

-stream static pressure, psf

psfa pounds per square foot absolute

psia pounds per square inch absolute

psid pounds per square inch 

differential

PCU pressure calibration unit q∞ free stream dynamic pressure, 

psf

Qflex gravity sensing servo accelerometer

Re Reynolds number based on 

reference wing chord

Re/ft unit Reynolds number

 s local semispan measured from 

the wing centerline, in.

S area, square inches (sq. in.) w S reference wing area, 

254.3553 sq. in. (also SAREA1)

t wing or canard thickness, in.

TC thermocouple T0 stagnation temperature,

o

F

V vertical tail volume, cubic 

inches (cu. in.)

W.L. model water line, in.

x local axial distance measured 

along the wing centerline chord 

from the wing apex, in.

XBAR1x-moment transfer distance 

measured in the body axis 

system from the BMC to MRC, 

positive in direction of positive 

thrust, 0.667 in.

y local semispan distance 

measured from the wing 

centerline, positive to the right, 

in.

YBAR1y-moment transfer distance 

measured in the body axis 

system from the BMC to MRC, 
4

positive in direction of positive 

side force, 0.0 in.

y span location of mean 

aerodynamic chord, in.

ZBAR1 z-moment transfer distance 

measured in the body axis 

system from the BMC to MRC, 

positive in direction of positive 

normal force, 0.984 in.

α angle of attack, degrees

β angle of sideslip, degrees

λ taper ratio

γ ratio of specific heats

ΛLE leading-edge sweep angle, deg

ΛTE trailing-edge sweep angle, deg

∆ denotes delta coefficient value;

in data repeatability analysis, 

∆’s are obtained by 

interpolating in each run to the 

nominal values of the 

independent variable, then 

averaging and subtracting the 

averages from the interpolated 

data

Subscripts

can canard

cham model chamber

cl centerline

lex leading-edge extension

r root

t tip

tw twin

w wing

Model Description and Test 

Apparatus

A generic fighter model featuring a 

65-degree cropped delta wing with sharp leading 

edges was used in this test. The model was 

designed and fabricated in the 1980’s for surface 

pressure and force and moment testing in 

subsonic, transonic, and supersonic wind tunnel

facilities in support of a multi-national Euler 

code validation program (reference 3) that 

focused on the prediction of vortex flow effects 

on slender wings at high angles of attack. The 

primary configurations that were tested in this 

program included the isolated wing with 

different leading-edge geometries (for example, 

round versus sharp) and a canard-wing 

arrangement. The model was subsequently used 

in a cooperative vortex flow aerodynamics 

research program in the 1980’s and 1990’s 

involving NASA Ames Research Center (NASA 

ARC), NASA LaRC, and the United States Air 

Force (USAF) Wright Aeronautical 

Laboratories. As part of this cooperative 

program, the model was tested in the NASA 

ARC 6- by 6-Foot Transonic/Supersonic Wind 

Tunnel (reference 4), NASA LaRC 7- by 10-

Foot HST (reference 5), NASA LaRC Unitary 

Plan Wind Tunnel (UPWT) (reference 6), and 

NASA LaRC 8-Foot TPT. Photographs of the 

model installed in the NASA LaRC 8-Foot TPT 

for Test 1057 discussed in this report are 

presented in figure 2. The wing had an NACA 

64A005 airfoil section from the 40 percent chord 

station to the trailing edge. A sharp 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 

pressure and balance instrumentation. In the 

isolated wing configuration, the fuselage tapered 

down to a small radius along approximately the 

forward 35-percent portion of its length, and it 

terminated 0.50 inches (model scale) from the 

apex of the wing. This portion of the fuselage 

could be replaced with an alternate forward 

fuselage section having an integral strut, or 

“gooseneck.” The model as designed in the 

1980’s included the installation of a 60-degree 

swept canard to the gooseneck to provide a 

closely-coupled canard-wing arrangement. The 

canard was co-planar with the wing and featured 

a biconvex circular-arc section with maximum 

thickness-to-chord ratio of 0.05. The canard 

area (left and right sides) was 16 percent of the 

reference wing area. Modifications were made 

to the model as part of the NASA and USAF 

cooperative research program to include a wing 

LEX mounted to the gooseneck in a co-planar

arrangement with the wing, a centerline vertical 

tail on the fuselage, and twin vertical tails

mounted to the wings. A flat-plate, 0.25-inch 
5

thick LEX having a 65-degree/90-degree

planform and symmetrically-beveled leading 

edges was fabricated for the investigation 

described in this report. The exposed area of the 

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

reference wing area. The LEX incorporated a 

pattern of 0.050-inch diameter through holes 

spaced 0.10 inch apart on center to provide a 

total porosity level (designated porosity level 1)

of 14.75 percent relative to the LEX exposed 

area. The fuselage and wings were modified 

with pockets to allow the installation of a 

centerline vertical tail or twin vertical tails with 

integral mounting pads. The primary model

components that were tested in the NASA LaRC 

8-Foot TPT were previously shown in figure 1

and figure 2. Planview and sideview sketches of 

the wing, LEX, canard, fuselage, and tails are 

provided in figure 3 through figure 6. Note that 

the bevels along the LEX leading and side edges 

were not porous because of constraints in the 

machining process that precluded drilling holes 

near the edges. In addition, a 0.625-inch wide 

strip along the centerline of the LEX was solid, 

since the LEX was bolted to the gooseneck in 

this region. Similarly, a 0.75-inch wide strip 

along each trailing edge of the LEX was solid 

where the LEX overlapped the wing leading 

edge in a tongue-and-groove arrangement. The 

same LEX was tested with 0 percent porosity 

(solid LEX) by applying sealing tape having 1.8 

mil thickness (0.0018 inches) along the lower 

surface to cover all of the through holes. Partial 

porosity was also possible by taping selected 

regions of the LEX lower surface. Four 

alternate levels of porosity were investigated in 

the 8-Foot TPT test. Porosity level 2 

corresponding to 9.4 percent of the LEX 

exposed area was obtained by applying tape to 

the triangular apex region as illustrated in 

figure 7. Conversely, applying tape to only the 

aft rectangular region yielded porosity level 3

(figure 7) corresponding to 5.4 percent of the 

LEX exposed area. Porosity level 4 representing 

6.1 percent of the LEX exposed area was 

obtained by applying tape to the lower surface 

except for a 0.575-inch wide strip along the LEX 

leading and side edges as shown in figure 7. 

Applying tape to only this 0.575-inch wide strip 

yielded porosity level 5, which corresponded to

8.7 percent of the LEX exposed area (figure 7). 

The size and positioning of the centerline and 

twin vertical tails were selected to provide 

approximately the same tail volume, which is 

the product of the total exposed tail area and the 

distance from the moment reference center 

(MRC) to the 25 percent mean aerodynamic 

chord location of the tail. Two sets of twin tails

were fabricated that featured, respectively, 

0-degree and 30-degree cant angles. Geometric 

details of the model are summarized in Table I.

The right wing upper surface was 

instrumented with a total of sixty four (64)

0.020-inch diameter pressure orifices distributed 

in three spanwise rows. Of the 64 available 

orifices, 47 were selected to populate a single

48-port ESP module. One orifice became

plugged during the initial wind-on check-out 

runs, so 46 orifices were used in Test 1057. 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, 

respectively). Sketches of the pressure orifice 

layout are shown in figure 3, figure 6, and 

figure 8, and the pressure orifice locations are 

listed in Table II. Cross sections of the wing at 

the three pressure measurement stations are 

sketched in figure 9. The wing upper surface 

static pressure distributions obtained in 

8-Foot TPT Test 1057 are compared to results 

obtained on the same model in a previous entry 

in the 7- by 10-Foot HST designated Test 202. 

The pressure orifice locations that were used in 

Test 202 differed slightly from those used in 

8-Foot TPT Test 1057 because of issues with 

certain orifices that were identified during the 

model installation leak-check process. The 

pressure orifice locations used in Test 202 are 

also defined in Table II. 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. In 8-Foot TPT Test 1057, there 
6

were 16 orifices at x c = 0.30, 16 orifices at 

x c = 0.60, and 14 orifices at x c = 0.80. Test 

202 in the 7- by 10-Foot HST featured 17 

orifices at x c = 0.30, 16 orifices at 

x c = 0.60, and 14 orifices at x c = 0.80. The 

distribution of taps in each row was intended to 

capture the pressure signature of a leading-edge 

vortex at high angles of attack. The pressures 

were measured using a single 48-port, 15 psid 

electronically-scanned pressure (ESP) module 

located inside the model as shown in figure 3. 

The 0.040-inch outer diameter (O.D.) stainless 

steel pressure lines inside the model were 

connected to the ESP module using urethane 

jumpers. One thermocouple wire was attached 

to the side of the ESP module to monitor the 

module surface temperature. The ESP 

electronics cable, reference and calibration 

pressure lines, and thermocouple wire were 

routed through channels machined into the 

internal balance strongback and then out along 

the model sting as illustrated in figure 10. The 

instrumentation bundle was then encased in 

protective plastic spiral wrap and fiberglass tape 

that extended along the model support system to 

an ESP interface box installed in the main 

support system arc sector base component. 

The model forces and moments were 

measured in Test 1057 using a NASA LaRC 

internally-mounted, six-component strain gage 

balance designated 755. A full balance 

calibration was performed prior to the entry in 

the 8-Foot TPT. The balance design loads and 

calibration accuracies are presented in Table III. 

The calibration accuracies expressed as 

aerodynamic coefficients based on the test 

conditions in Test 1057 are presented in 

Table IV. Sketches of the balance inside the 

model were previously shown in figure 4 and 

figure 5. The balance wiring was routed 

internally to the sting. The balance moment 

center (BMC) was located at model station 

(M.S.) 21.811 which corresponded to 

approximately 59.8 percent of the distance along 

the wing centerline chord measured from the 

wing apex ( x c = 0.598). The MRC was taken 

about the 57 percent centerline chord location 

( x c = 0.57) or M. S. 21.144 as shown in 

figure 3 through figure 5. The force and 

moment coefficients for all configurations were 

based on the reference wing area, Sw. The 

NASA LaRC balance 842A was used in 

7- by 10-Foot HST Test 202. The corresponding 

balance design loads and calibration accuracies 

are also shown in Table III. 

The model base area was negligible, 

consequently, base pressures were not measured 

in 8-Foot TPT Test 1057. Two 0.040-inch O.D. 

stainless steel tubes were run externally and 

diagonally opposed along the sting and extended 

inside the model fuselage cavity. One chamber 

pressure tube terminated approximately 0.5 

inches aft of the balance-to-sting draw nut, and 

the second chamber pressure tube terminated 2.0 

inches forward of the fouling strip as shown in 

figure 10. The tubes were connected to 

individual 2.5 psid pressure tranducers located in 

the plenum region surrounding the test section. 

Full calibrations of the pressure transducers 

were conducted during the model installation 

process. 

The NASA LaRC sting number 20 served a 

dual role of providing an internal passageway 

for the balance wiring bundle and adapting the 

model and balance assembly to the tunnel 

support hardware. The sting was a taper fit to 

the 8-Foot TPT motorized yaw coupler. The yaw 

coupler had a 7.5-degree angle of attack offset 

and provided yaw angle variations from 

approximately -7 degrees to +7 degrees prior to 

sting deflections due to aerodynamic loads. The 

yaw coupler was also calibrated during the 

model installation process. Additional taper fit 

joints in the model support system downstream 

of the yaw coupling included a split knuckle 

with 5-degree pitch angle offset, 8-Foot TPT 

knuckle adapter, and 12-inch offset adapter. 

This multiple-joint assembly was installed to the 

main support system base and motor-driven arc 

sector which featured a variable-rate angle of 

attack drive. The arc sector could be moved 

axially about 6 feet, and it was translated to 

position the tip of the model LEX at a fixed 

tunnel station of 79 inches measured from the 

origin of the slots in the floor and ceiling of the 
7

transonic test section. This positioned the model 

at a desirable location with respect to the LVS 

flow visualization system. 

Transition grit was not applied to the model, 

since the flow was assumed to separate at the 

sharp leading edges at all conditions of interest 

in the current experiment. In addition, a suitable 

gritting strategy based on the criteria in 

reference 7 to cause transition of the boundary 

layer associated with vortex-induced reattached 

flow on the wing upper surface has not been 

established. Reference 8 summarizes many of 

the challenges associated with transition grit 

applications for high angle-of-attack 

experimentation.

A fouling strip circuit was installed on the 

sting near the model base (figure 10). Model-tosting fouling was not encountered at any of the

wind-on conditions in Test 1057.

Wind Tunnel Facilities and Test 

Conditions

The investigation was conducted in the 

NASA LaRC 8-Foot TPT and was designated 

Test 1057. Photographs of the general research

fighter model installed in the 8-Foot TPT test 

section were previously shown in figure 2. This 

wind tunnel facility was designed for operation 

as a continuous-flow, closed-return, variablepressure wind tunnel with control capability to 

independently vary Mach number, stagnation 

pressure, stagnation temperature, and humidity. 

The test section was square with corner fillets 

with 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; the slot-width contour provided a 

gradient-free test section 50 inches long for 

Mach numbers equal to or greater than 1, and 

100 inches long for Mach numbers less than 1. 

The stagnation pressure could be varied from 

0.25 to 2 atmospheres. Reference 9 provides a 

detailed description of the 8-Foot TPT. Note 

that this facility was permanently closed in 

1995.

The majority of Test 1057 was performed at 

Mach = 0.50, 0.85, and 1.20 at a unit Reynolds 

number, Re/ft, of 2.5 million (Re = 4.92 million 

based on a wing reference chord of 23.622 

inches) and a total temperature of 120 degrees 

Fahrenheit. Subsets of runs were performed at 

Mach = 0.50 and unit Reynolds numbers of 1.5 

and 3.5 million (2.95 million and 6.89 million 

based on the wing reference chord). The test 

conditions for Test 1057 are summarized in 

Table V. The tolerances for the tunnel condition 

setpoints for the Mach number, total pressure, 

and total temperature were typically +/-0.002, 

+/-5 pounds per square foot absolute (psfa), and 

+/-2 degrees Fahrenheit (o

F), respectively. The 

tunnel air dew point was maintained at sufficient 

levels to minimize water vapor condensation 

effects during all phases of the test except during 

the LVS flow visualization runs.

Previous test results obtained at subsonic 

speeds on the same model in a prior entry in the 

NASA LaRC 7- by 10-Foot HST are included in 

the section on Tunnel-to-Tunnel Data 

Reproducibility. This investigation was

designated Test 202. The NASA LaRC 

7- by 10-Foot HST was a continuous-flow 

subsonic-transonic atmospheric facility. In the 

closed test section configuration, the speed range 

was from approximately Mach = 0.06 to 

Mach = 0.94 depending on the model size. The 

test section was approximately 6.584 feet high 

by 9.574 ft wide, and the mean cross-sectional 

area of the test section for use in blockage 

calculations was 62.1256 square feet. The tunnel 

operated at ambient temperature and pressure 

and continuously exchanged air with the 

surrounding atmosphere. Test results were 

obtained at Mach = 0.50 in both facilities, 

which enabled a tunnel-to-tunnel comparison of 

the wing surface static pressure distributions and 

the six-component force and moment data for 

several configurations. The model installed on 

the high angle-of-attack static stability support 

system in the 7- by 10-Foot HST test section is 

shown in the photographs in figure 11. The test 

conditions at Mach = 0.50 in the NASA LaRC 

7- by 10-Foot HST corresponded to a unit 

Reynolds number of approximately 3.0 million
8

(5.91 million based on the wing reference 

chord), and free-stream dynamic pressure of 

317 pounds per square foot. Reference 10

provides a detailed description of the calibration, 

operation, and testing capabilities of the NASA 

LaRC 7- by 10-Foot HST. This facility was 

permanently closed in 1994.

The model angle of attack and angle of 

sideslip were determined via appropriate Euler 

angle transformations using the output from an 

accelerometer mounted in the base of the 

8-Foot TPT arc sector support system, the 

output from a potentiometer installed in the 

motorized yaw coupler, balance-to-support 

system and balance-to-model misalignment 

angles, and corrections applied to account for 

aeroelastic or mechanical deflections of the 

model, balance, and support system assembly 

due to aerodynamic loads. The tolerances for the 

angle of attack and sideslip setpoints were 

typically +/-0.1 degrees and +/-0.05 degrees, 

respectively. 

The model was painted flat black for the

LVS flow visualization. In this phase, the 

dewpoint was allowed to vary in order to 

promote condensation in the test section. LVS 

flow visualization was conducted on the model 

with solid LEX, porous LEX, and isolated wing

with the centerline vertical tail at Mach = 0.85 

only, since this was the test condition of primary 

interest in Test 1057. Visualization of the vortex 

flows at Mach = 0.50 in the 8-Foot TPT was not 

performed because of the prohibitively large 

amounts of water injection into the tunnel circuit 

necessary to promote local condensation. 

Video images of the LVS cross-flow patterns

were obtained at selected angles of attack and 

model stations using a miniature color video 

camera mounted to the non-rotating base section 

of the motorized yaw coupler. The LVS images 

were useful in the interpretation of the ESP 

upper surface static pressure distributions and 

the aerodynamic force and moment coefficient 

trends. A more detailed description of the 

8-Foot TPT LVS system is provided in the 

Experimental Techniques section and in 

reference 5.

ESP and six-component force and moment 

measurements were simultaneously obtained at 

Mach = 0.50, 0.85, and 1.20 for all 

configurations presented in this report. Previous 

testing experience with this model in the NASA 

ARC 6- by 6-Foot Transonic/Supersonic Tunnel 

(reference 4) and the NASA LaRC 

7- by 10-Foot HST indicated that sufficient slack 

could be provided in the ESP wiring and tubing 

bundle extending from the routing channels in

the balance strongback to the sting to mitigate 

the effects of bridging the balance on the force 

and moment measurements. The lowest angle of 

attack that could be obtained was approximately 

+1.5 degrees because of the pitch angle offsets 

in the model support system components. The 

nominal angle of attack range was +1.5 degrees, 

+2 degrees to +24 degrees in 2-degree 

increments, and to the maximum angle of attack 

that was possible for a given configuration and 

Mach number. Sideslip sweeps were also 

conducted at angles of attack of 8, 12, 16, 20, 

and 24 degrees for selected configurations and 

Mach numbers. The sideslip angle was varied 

from 0 degrees to 7 degrees in 1-degree 

increments and then to maximum possible 

sideslip angle (typically less than 8 degrees). 

The sideslip angle was returned to 0 degrees, 

and the sweep was completed by varying the 

sideslip angle from 0 to -7 degrees in 1-degree 

increments and then to the maximum negative 

sideslip angles (typically less than -8 degrees). 

This sequencing of the sideslip angle was 

intended to mitigate the potential effects of 

aerodynamic hysteresis that have been observed 

on vortex- and shock-dominated flow fields in 

previous investigations (reference 4). The angle 

of attack and sideslip sweeps were performed in 

pause modes. The Experimental Techniques 

section describes the ESP and force and moment 

measurement techniques in more detail.

Experimental Techniques

Laser Vapor Screen Technique

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, 
9

transonic, and supersonic speeds. Water is 

injected in sufficient quantity into the tunnel 

circuit, typically downstream of the supersonic 

nozzle or the diffuser section, to cause 

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 in a 

plane perpendicular to the longitudinal axis of 

the tunnel or to the body axis of the model. 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. At supersonic 

speeds, the temperature drop from the expansion 

in the supersonic nozzle causes the water vapor 

to condense into a fine fog. Since condensation 

first appears in the free stream at supersonic 

speeds, the vortex flows appear as dark regions 

in the absence of scattered light surrounded by a 

light background. At subsonic and transonic 

speeds, condensed water vapor generally first 

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 light-scattering patterns often occurs at 

transonic speeds.

A fiber-optic-based laser vapor screen system 

was established in the 8-Foot TPT in 1990 to 

visualize the vortex-dominated flow fields about 

small-scale models of fighter aircraft, 

commercial transport airplanes, and missiles at 

subsonic and transonic speeds. A detailed 

description of this system is provided in 

reference 5. A fiber optic cable delivered a 

beam from an argon laser located outside the 

tunnel plenum to a light sheet optics package 

located in the ceiling of the test section. The 

fiber-optic-based beam delivery system 

contained five principal components: laser-tofiber coupler, armored fiber optic cable, remote 

light sheet generator, rotating mirror, and optics 

motor controller. The system was designed to 

be used with virtually any argon-ion laser 

system operating in either continuous wave or 

multimode with beam diameters from 

approximately 0.0315 inches to 0.0709 inches. 

The light sheet optics package allowed variation 

of the sheet thickness, divergence or spread of 

the light sheet, and sheet rotation relative to the 

model. A 3- by 3-inch mirror mounted onto a 

rotational stage with 360 degrees of continuous 

rotation directed the light sheet through the 

optical window and to the desired station on the 

model. The light sheet was aligned to be 

perpendicular to the model surface at 

approximately the 50 percent wing chord station

and an angle of attack of 16 degrees. Because 

the light sheet swept in an arc along the model, it 

was non-orthogonal with respect to the model 

surface at all other conditions.

Water was injected into the tunnel circuit 

from a 150-gallon tank of deionized water 

located in a room on top of the plenum shell to 

an array of six atomizer nozzles installed in the 

ceiling region of the diffuser section. The 

amount of water that was injected into the tunnel 

was remotely regulated from the wind tunnel 

control room via a solenoid switch and a onehorsepower pump positioned on the discharge 

side of the water tank.

Documentation of the LVS images was 

obtained using a miniature video camera with 

360 television lines of horizontal resolution and 

a fixed focal length lens contained in a 

cylindrical housing mounted onto the model 

support system. This camera provided a 

perspective aft of the model which looked 

upstream along the model centerline. The video 

image perspective remained constant throughout 

the pitch angle range because there was no 

relative motion between the camera and the 

model. The video image perspective did vary, 

however, if the model yaw angle was changed. 

Selected portions of the videotape recording 

were digitized and converted to audio video 

interleave (AVI) format using a video frame 

grabber installed in a personal computer. The 

frame grabber software allowed the precise 

capture and enhancement of the individual LVS 

frames that appear in this report.
10

ESP Measurements 

Surface static pressure measurements were 

obtained in Test 1057 at 46 discrete locations on 

the model using a single, internally-mounted 48-

port ESP module. One pressure orifice became 

non-functional shortly after wind-on operations 

were initiated. The remaining port was 

dedicated to a reference pressure line. The 

8-Foot TPT ESP data acquisition system was 

designated System 780B and was interfaced 

with the wind tunnel data acquisition system. 

The System 780B accommodated external 

modules, modules mounted internally to the 

model, or combinations of internal and external 

modules. The internal volume of the 65-degree 

cropped delta wing model was sufficient to 

contain a single 48-port, 15 psid ESP model. A 

15 psi pressure calibration unit (PCU) was used 

in this experiment with a digitally-controlled 

pneumatic source that provided valve control 

and generated calibration pressures for the ESP 

scanner. The ESP module pressure 

measurement accuracy was assumed to be 

+/-0.05% full-scale (F. S.). The ESP module 

pressure range was selected on the basis of prior 

subsonic testing of this model with solid LEX 

(reference 4). The uncertainties for this module 

expressed in terms of the static pressure 

coefficient for the test conditions that were run 

in the 8-Foot TPT Test 1057 are listed in 

Table VI. The wing upper surface static 

pressure coefficients obtained in Test 1057 are 

compared in this report to the vacuum pressure 

coefficient, Cp,v, and the pressure coefficient 

corresponding to the local speed of sound, Cp

*

. 

The corresponding values of Cp,v and Cp

*

 at 

Mach = 0.50, 0.85, and 1.20 are listed in 

Table VII. The pressure distributions from 

7- by 10-Foot HST Test 202 were obtained with 

a similar System 780B data acquisition unit with 

an internal 15 psid ESP module having the same 

calibration accuracies as those quoted for the 

8-Foot TPT testing. In the tunnel-to-tunnel 

comparisons at Mach = 0.50, the uncertainties 

for the ESP modules expressed as static pressure 

coefficients are based on the 8-Foot TPT test 

conditions at this Mach number. 

 The ESP data acquisition rate during the 

testing was 10 frames per second for 2 seconds, 

for a total of 20 frames per data point. A dwell 

time of 5 seconds was specified to allow the 

pressures to stabilize before acquiring a data 

point. 

The standard ESP calibration consisted of 

five points that were used to determine a quartic 

polynomial representation of the pressurevoltage signature of each transducer or sensor. 

Full wind-on calibrations were performed prior 

to each run series once the ESP module 

temperature had stabilized. Updated calibrations 

were also performed for each change in the freestream Mach number. 

Strain Gage Balance Measurements 

Force and moment data were obtained with a 

6-component electrical strain gage balance 

designated NASA LaRC 755. The 755 balance 

design loads and the balance calibration 

accuracies (95% confidence level) expressed in 

percent full-scale (% F. S.) were previously 

shown in Table III. The balance accuracies were 

converted to microvolts (V) and, also, to 

pounds (lbs) or inch-pounds (in-lbs) as shown in 

Table III. The latter values were used to 

estimate the measurement accuracies expressed 

as aerodynamic force and moment coefficients 

for the conditions in Test 1057 as shown in 

Table IV. 

The balance data acquisition rate during the 

testing was 30 frames per second for 2 seconds, 

for a total of 60 frames per data point. All force 

and moment data were acquired in a pitch-pause 

mode. A dwell time of 5 seconds at each 

setpoint was dictated by the simultaneous 

acquisition of the ESP data. 

Force and moment and surface pressure 

measurements were simultaneously obtained 

despite the bridging of the balance with the onboard ESP cable, reference and calibration 

pressure lines, and thermocouple wire. 

Precautions were taken to mitigate any bridging 

effects on the balance force and moment 

measurements by distributing the ESP 
11

instrumentation bundle through the routing 

channels in the balance strongback and 

providing a flexible bridge from the model to the 

sting. Calibrated weights were placed at known 

locations on the model to apply prescribed 

values of the normal force, pitching moment, 

and rolling moment. A hand-held force gage 

was also used to apply check loads to all six 

force and moment components. The applied 

loads were compared to the computed loads 

from the wind tunnel data system, which 

indicated that simultaneous acquisition of the 

balance and ESP measurements was a valid 

testing approach in the current application. 

Model Chamber Pressure Measurements

Model chamber pressures were measured at 

two diagonally-opposed locations. One chamber 

pressure tube terminated approximately 

0.5 inches aft of the balance-to-sting draw nut, 

and the second chamber pressure tube 

terminated 2.0 inches forward of the fouling 

strip as previously shown in figure 10. The 

0.040-inch O.D. stainless steel pressure tubes 

were run along the sting and model support 

system and routed to the plenum region 

surrounding the test section where they 

connected to individual 2.5 psid pressure 

transducers. The manufacturer-specified 

measurement accuracy (95% confidence level) 

as a percent of full-scale is +/-0.1% for these 

transducers (+/-0.0025 psid). 

Pitch Angle Measurements 

The primary type of instrumentation in use at 

the 8-Foot TPT for pitch angle measurement was 

a gravity-sensing servo accelerometer (Qflex). 

Direct and indirect methods of model attitude 

measurement were used. The direct 

measurement used an accelerometer mounted in 

the model. The indirect measurement featured 

an accelerometer installed in the main arc sector 

support system with corrections applied to 

account for aeroelastic or mechanical deflections 

of the model, balance, sting, and support system 

component assembly. There was insufficient 

internal volume in the 65-degree cropped delta 

wing model to accommodate a Qflex package. 

Consequently, the indirect method of attitude 

measurement was used in Test 1057. For static 

(un-accelerated) model conditions, the Qflex 

measures changes in angle relative to the 

horizontal by determining the differences in the 

component of the force due to gravity acting 

parallel to its sensitive axis. Although the 

instrument response to acceleration is linear, its 

response to changes in attitude relative to the 

local gravity vector in un-accelerated conditions 

is sinusoidal. A Qflex calibration was 

performed by installing a digital inclinometer to 

the precision-machined LEX upper surface. A 

14-point calibration was performed at pitch 

angles from approximately +1.5 degrees to a 

maximum static pitch angle of +25.3 degrees. 

The standard deviation from the Qflex 

calibration was approximately 0.005 degrees. 

Since models in the 8-Foot TPT were normally 

forward of the arc sector center of rotation, the 

model nose could approach the tunnel ceiling as 

the model rotated to high angles of attack. 

Measurements were taken from the test section 

ceiling to the apex of the model LEX at the 

highest pitch angles during the Qflex calibration. 

At a pitch angle of 24 degrees, the LEX apex 

was 19.125 inches from the ceiling; this distance 

diminished to 16.875 inches at the maximum 

static pitch angle of 25.3 degrees. 

Corrections to account for aeroelastic or 

mechanical deflections of the model, balance, 

and sting assembly due to aerodynamic loads 

were based on in-tunnel sting and balance 

deflection calibrations. Deflections due to 

normal force, pitching moment, rolling moment, 

side force, and yawing moment were obtained 

using calibrated weights suspended on a pan and 

attached to the balance calibration fixture via a 

double knife-edge assembly. Prescribed loads 

were applied at predetermined locations relative 

to the BMC, and the corresponding deflections 

were recorded using a digital inclinometer 

installed on the balance calibration fixture and 

referenced to the Qflex in the main support 

system arc sector. A second inclinometer was 

installed at an intermediate location on the 

support system to check for mechanical 

misalignment. The sting deflection calibrations 

also provided an opportunity to check the data 
12

acquisition and data reduction system by 

comparing the applied loads to the computed 

loads. 

Yaw Angle Measurements

The 8-Foot TPT motorized yaw coupler 

provided the primary measurement for the 

model yaw angle. Output from the yaw coupler 

was obtained using an onboard potentiometer. 

The yaw coupling was calibrated using a digital 

inclinometer mounted to the balance calibration 

fixture with the yaw coupler temporarily rolled 

90 degrees. This calibration was coordinated 

with the sting deflection calibrations for side 

force and yawing moment.

Corrections

Flow angularity

Several previous tests of three-dimensional 

(3-D) models in the 8-Foot TPT indicated that 

flow angularity in the test section was small and 

was typically less than approximately 0.05 

degrees. Consequently, flow angle runs were 

not performed in Test 1057, and the force and 

moment data were not corrected for the assumed 

small effects of tunnel flow angularity. Similar 

assumptions were made for the data obtained in 

the 7- by 10-Foot HST.

Other corrections

Following standard procedures at the 

8-Foot TPT, wall interference, buoyancy, and 

blockage corrections were not applied to the data 

because of the slotted test section and the small 

size of the model (reference 9). Jet boundary 

and blockage corrections were applied to the test 

data from the 7- by 10-Foot HST that are 

presented in this report according to the 

procedures in references 11 and 12.

Chamber pressures

The model chamber pressure measurements 

were used to correct the balance axial force to a 

condition of free-stream static pressure at the 

model base. Base pressure corrections were not 

applied, since the model base area was 

essentially zero.

Discussion of Results

Data plots created using data engineering 

scripting language (DESL) scripts (reference 13) 

are presented in this section that represent an 

extensive sampling of the results obtained in 

8-Foot TPT Test 1057. Several basic plot 

formats are used to illustrate the upper surface 

static pressure coefficient distributions and the 

six-component aerodynamic force and moment 

coefficient trends in angle-of-attack sweeps and 

sideslip sweeps. The wing pressure distributions 

are typically plotted in the three spanwise rows 

superimposed onto an isometric view of the 

right-hand wing. The pressure distributions are 

plotted in this format for selected angles of 

attack, angles of sideslip, Mach numbers, 

Reynolds numbers, model configurations, and 

within-test and tunnel-to-tunnel repeat runs. 

More conventional two-dimensional plot 

formats are used to extend the analysis of the 

within-test data repeatability by displaying the 

delta pressure coefficient values (∆Cp)

corresponding to the differences in the pressure 

coefficients in repeat runs as a function of the 

pressure orifice numbers 1 through 47, which 

progressed from the inboard to the outboard 

orifice in each spanwise row beginning with the 

first pressure measurement station at x/c = 0.30. 

The data obtained in the three spanwise rows are 

rendered more distinguishable by plotting them

in separate groups as a function of the wing 

orifice number. The upper and lower bounds 

corresponding to the ESP module calibration 

pressure accuracies from Table VI are 

superimposed on the delta pressure coefficient 

values to better quantify the data scatter. Similar 

delta pressure coefficient plots are presented in 

the analysis of tunnel-to-tunnel data 

reproducibility, where the differences obtained 

in the pressure measurements at common 

orifices in the three spanwise rows are 

displayed. A typical force and moment data 

plot format used in this report corresponds to 

angle of attack sweeps at a nominal sideslip 

angle, β, of 0 degrees, where the longitudinal 

aerodynamic coefficients in the stability axis 

system are displayed (CL versus angle of attack

(α), CL versus CD, and CL versus Cm). The 

corresponding lateral-directional aerodynamic 
13

coefficients in the body axis system (Cn, Cl, and 

CY) are plotted versus α. Another plot format 

corresponds to sideslip sweeps, where the 

longitudinal aerodynamic coefficients in the 

stability axis system and the lateral-directional 

aerodynamic coefficients in the body axis 

system are plotted as functions of β at selected 

fixed angles of attack. Within-test and tunnelto-tunnel data repeatability assessment is based 

largely on delta aerodynamic coefficient plots, 

which show each of the six aerodynamic force 

and moment coefficients in the body axis system 

plotted versus α along with delta coefficient 

plots (∆CN, ∆CA, ∆Cm, ∆Cn, ∆Cl, ∆CY) to show 

the variability in the data compared to the 

95 percent confidence limits for the 755 balance 

accuracies derived from the balance calibration. 

The delta coefficients were obtained by 

interpolating in each angle-of-attack sweep to 

the nominal values of the independent variable, 

then averaging and subtracting the averages 

from the interpolated data. The initial and repeat 

runs in 8-Foot TPT Test 1057 were not 

consecutive but, instead, separated by several 

angle-of-attack and/or sideslip sweeps at the 

same and different Mach numbers and spanned

several tunnel operating shifts. The 

tunnel-to-tunnel data comparisons at 

Mach = 0.50 offered a more robust assessment 

of data reproducibility, since the 8-Foot TPT and 

7- by 10-Foot HST tests were performed over a 

period of two years. Wherever possible, 

common plot scales were used. However, in 

many instances the plot scales were adjusted to 

adequately display the results. This could 

obscure the 95 percent confidence limits for the 

ESP and balance accuracies in the delta 

coefficient plots, since these limits were often so 

small compared to the data scatter that they 

could not be distinguished from the horizontal 

axis. 

The data plots are presented in seven broad

sections, each containing the wing surface 

pressure distributions and the six-component 

force and moment measurements. These 

sections include (1) Reynolds number 

assessment at Mach = 0.50 for the solid and 

porous LEX configurations; (2) 8-Foot TPT 

within-test data repeatability at Mach = 0.5 and 

0.85; (3) 8-Foot TPT and 7- by 10-Foot HST 

tunnel-to-tunnel data comparisons at 

Mach = 0.50; (4) comparisons of the solid LEX 

to porous LEX with full porosity at 

Mach = 0.50, 0.85, and 1.20 (LVS flow 

visualization images obtained at Mach = 0.85 

are included); (5) comparisons of full porosity 

versus partial porosity obtained by partitioning 

the porous regions on the LEX at Mach = 0.50 

and 0.85; (6) comparisons of the vertical tail 

arrangements with solid LEX, including the 

centerline tail, twin uncanted tails, twin canted 

tails, and tails off at Mach = 0.50 and 0.85; and 

(7) comparisons of the solid LEX and 

closely-coupled canard at Mach = 0.50 and 0.85.

Reynolds Number Assessment

Pressure Distributions

Figure 12 and figure 13 show the wing upper 

surface static pressure distributions obtained 

with the solid LEX and the porous LEX, 

respectively, at Mach = 0.50 and unit Reynolds 

numbers, Re/ft, of 1.5 million, 2.5 million, and 

3.5 million. Both configurations featured the 

centerline vertical tail. Each figure contains data 

obtained at angles of attack of 10 degrees and 

20 degrees. There is some variability in the data 

(see figure 13(b) at α = 20 degrees and 

x/c = 0.80, for example). However, no 

significant effects of the Reynolds number are 

apparent in the pressure distributions within the 

range of Reynolds number tested and to the plot 

scales shown in figure 12 and figure 13.

The data presented in figure 12

corresponding to the solid LEX were used to 

create the delta pressure coefficient plots in 

figure 14 and figure 15. The differences in the 

pressure coefficients at Re/ft = 1.5 million and 

Re/ft = 3.5 million and α = 10 and 20 degrees 

are shown in figure 14, and the corresponding 

pressure coefficient differences at Re/ft = 2.5 

million and Re/ft = 3.5 million are presented in 

figure 15. The ESP module pressure 

measurement accuracies representing the upper 

and lower 95 percent confidence limits at 

Mach = 0.50 are displayed as horizontal dashed 

lines in the delta pressure coefficient plots. The 
14

delta coefficients in figure 14 are generally 

outside the pressure measurement accuracy 

limits. Modified upper and lower limits 

corresponding to approximately 0.5 percent F. S. 

range of the ESP module (compared to the 

assumed pressure measurement accuracy of 

0.05 percent F.S.) would effectively bound a 

majority of the delta coefficient values at all 

pressure measurement stations and both angles 

of attack. The delta coefficients in figure 15 are 

also typically outside the assumed pressure 

measurement accuracy bounds, although the 

data scatter is reduced by nearly 50 percent

compared to the results shown in figure 14.

The data presented in figure 13

corresponding to the porous LEX were used to 

create the delta pressure coefficient plots in 

figure 16 and figure 17. The differences in the 

pressure coefficients at Re/ft = 1.5 million and 

Re/ft = 3.5 million and α = 10 and 20 degrees 

are shown in figure 16, and the corresponding 

pressure coefficient differences at 

Re/ft = 2.5 million and Re/ft = 3.5 million are 

presented in figure 17. The delta coefficients in 

figure 16 are outside the pressure measurement 

accuracy limits. The data scatter is increased 

compared to the corresponding results obtained 

with the solid LEX, and modified upper and 

lower limits corresponding to approximately 

1 percent F. S. range of the ESP module 

(compared to the assumed pressure measurement 

accuracy of 0.05 percent F.S.) would effectively 

bound a majority of the delta coefficient values 

at the three pressure measurement stations and at 

both angles of attack. The delta coefficients in 

figure 17 are also typically outside the assumed 

pressure measurement accuracy bounds, 

although the data scatter is again reduced by 

nearly 50 percent compared to the results shown 

in figure 16. The increased data scatter with the 

porous LEX may be caused by unsteady and 

Reynolds number-sensitive flow through the 

0.050-inch-diameter holes in the LEX. 

However, the scatter in the delta pressure 

coefficients that is observed in figure 14 through 

figure 17 at Re/ft = 1.5 to 3.5 million is similar 

to that observed in repeat runs at the same 

Reynolds number as shown in the section 

Within-Test Data Repeatability.

Force and Moment Measurements

Figure 18 and figure 19 show the 

longitudinal aerodynamic coefficients obtained 

in angle of attack sweeps at a sideslip angle of 

0 degrees with the solid LEX and the porous 

LEX, respectively, at Mach = 0.50 and unit 

Reynolds numbers of 1.5 million, 2.5 million, 

and 3.5 million. The maximum angle of attack 

increased with increasing Reynolds number 

because of the higher dynamic pressures and 

correspondingly higher sting deflections due to 

aerodynamic loads. The lift, drag, and pitching 

moment coefficients do not exhibit any

discernible sensitivity to the Reynolds number 

within the range of Reynolds number tested and 

to the plot scales shown in figure 18 and 

figure 19.

Figure 20 and figure 21 show the delta 

coefficient plots of the normal force, axial force, 

and pitching moment coefficients (∆CN, ∆CA, 

∆Cm) versus α corresponding to the solid LEX 

and porous LEX, respectively. The 95 percent

confidence limits for the 755 balance accuracies 

derived from the balance calibration expressed 

in terms of the aerodynamic coefficients at 

Mach = 0.50 are depicted as horizontal dashed 

lines in the delta coefficient plots. Computation 

of the delta aerodynamic coefficients was 

constrained to angles of attack from 2 degrees to 

26 degrees for the solid LEX and from 2 degrees 

to 24 degrees for the porous LEX to avoid 

extrapolation at the higher angles of attack. The 

range of angle of attack within which the delta 

coefficients are reasonably bounded by the 

balance calibration accuracy limits depends on 

the aerodynamic coefficient and the LEX 

configuration. The balance calibration accuracy 

limits for normal force, axial force, and pitching 

moment are 0.05 percent F. S., 0.26 percent 

F. S., and 0.09 percent F. S., respectively 

(Table III). With the solid LEX (figure 20), the 

scatter in the normal force coefficient is 

generally bounded by the balance accuracy 

limits from α = 2 to 16 degrees; the scatter in the 

axial force coefficient is typically bounded up to 
15

α = 22 degrees; and pitching moment coefficient

scatter is bracketed up to α = 10 degrees.

Expressed in terms of modified balance 

accuracy limits, the ∆CN values do not exceed 

approximately 0.2 percent of the balance normal 

force design load; the ∆CA values are within 

0.44 percent of the balance axial force design 

load; and the ∆Cm values are bounded by limits 

corresponding to 0.59 percent of the balance 

pitching moment design load. The normal force 

coefficient data scatter with the porous LEX 

(figure 21) exceeds the balance accuracy limits 

at all angles of attack. However, the ∆CN values 

are effectively bracketed by modified limits 

corresponding to 0.22 percent of the balance 

design load. The scatter in the axial force 

coefficient is effectively contained within the 

balance accuracy limits at all angles of attack 

from 2 degrees to 24 degrees. The pitching 

moment coefficient data scatter is bounded by 

the balance accuracy limits up to α = 18 degrees, 

and the ∆Cm values are bounded by modified 

limits corresponding to 0.3 percent of the 

balance design load at the higher angles of 

attack. 

The analysis of the surface pressure 

distributions and the longitudinal aerodynamic 

force and moment coefficients does not reveal a 

sensitivity that could be directly attributed to 

changes in the Reynolds number from 

Re/ft = 1.5 million to Re/ft = 3.5 million. The 

observed data scatter is similar to results 

obtained in testing of slender wing models at 

high angles of attack and at subsonic through 

supersonic speeds (reference 4 and reference 6, 

for example). In addition, the delta coefficient 

values do not exceed the data scatter observed in 

the current test in repeat runs obtained on 

selected configurations at the same Reynolds 

number as discussed in the next section WithinTest Data Repeatability. Consequently, the 

selection of Re/ft = 2.5 million for the 

remainder of Test 1057 was considered 

reasonable in terms of aerodynamic data validity 

and wind tunnel facility power requirements at 

the subsonic and transonic Mach number 

conditions. 

Within-Test Data Repeatability

Pressure Distributions

Figure 22 and figure 23 show the wing upper 

surface static pressure distributions obtained in 

repeat runs with the solid LEX and centerline 

vertical tail at Mach = 0.50 and 0.85, 

respectively, and a unit Reynolds number of 

2.5 million. Each figure contains data obtained 

at α = 12, 16, 20, and 24 degrees. Similar 

results obtained with the porous LEX are 

presented in figure 24 and figure 25

corresponding to Mach = 0.50 and 0.85, 

respectively. For each configuration, the 

acquisition of the repeat runs spanned at least 

two operating shifts without intervening model 

changes. There is some variability in the data, 

particularly with the porous LEX at the higher 

angles of attack (figure 24 and figure 25). 

However, a qualitative assessment of the results 

presented in figure 22 through figure 25 suggests 

the within-test repeatability of the wing upper 

surface static pressure distributions is 

reasonable, particularly given the presence of 

highly-separated flows at both Mach numbers 

and the assumed development of shock waves 

co-existing with the leading-edge vortices at 

Mach = 0.85 (reference 3 and reference 4).

The data presented in figure 22 and figure 23

corresponding to the solid LEX at Mach = 0.50 

and 0.85, respectively, were used to create the 

delta pressure coefficient plots in figure 26 and 

figure 27. The ESP module pressure 

measurement accuracies representing the upper 

and lower 95 percent confidence limits at 

Mach = 0.50 and 0.85 are displayed as 

horizontal dashed lines in the delta pressure 

coefficient plots. In addition, the pressure 

coefficient corresponding to the local speed of 

sound, Cp*, is depicted as a horizontal dashed 

line in selected plots of the upper surface static 

pressure coefficient distributions obtained in the 

repeat runs. At Mach = 0.50 and α = 12 and 

16 degrees (figure 26(a) and figure 26(b)), the 

differences in the static pressure coefficients 

between the repeat runs are bounded reasonably 

well by the ESP module pressure measurement 

accuracies. Larger delta coefficient values are 
16

observed at α = 20 and 24 degrees (figure 26(c)

and figure 26(d)), although the maximum ∆Cp

values do not exceed approximately 0.5 percent

of the F. S. range of the ESP module. The 

pressure coefficients are subcritical at all angles 

of attack and pressure measurement stations, 

although the suction pressure peak induced by 

the vortex flows at α = 24 degrees and 

x/c = 0.80 approaches the sonic condition. The 

data scatter between repeat runs at Mach = 0.85 

(figure 27) also increases as the angle of attack 

increases. Modified pressure measurement 

accuracy limits corresponding to approximately 

+/-0.8 percent of the F. S. range of the ESP 

module would effectively bound the majority of 

∆Cp values up to α = 24 degrees. There is an 

apparent bias in the delta coefficient plots such 

that the ∆Cp values are almost exclusively 

positive at α = 16, 20, and 24 degrees. These 

results may be indicative of a thermal sensitivity 

of the ESP module pressure transducers and an 

insufficient frequency of on-line ESP module 

calibrations to compensate for thermal effects at 

the transonic speeds. The majority of pressure 

coefficients at α = 12 degrees exceed the critical 

value (figure 27(a)), and all pressure coefficients 

are supercritical (corresponding to locally 

supersonic flow) at the higher angles of attack 

(figure 27(b) through figure 27(d)).

The data presented in figure 24 and figure 25

corresponding to the porous LEX at 

Mach = 0.50 and 0.85, respectively, were used 

to create the delta pressure coefficient plots in 

figure 28 and figure 29. The differences in the 

pressure coefficients obtained in repeat runs at 

Mach = 0.50 (figure 28) are significantly larger 

with the porous LEX compared to the results 

previously shown for the solid LEX in figure 26. 

There is no indication of systematic differences 

between the repeat runs, that is, the data scatter 

appears to be randomly-distributed. The largest 

∆Cp values typically occur in regions of vortexinduced flow reattachment at x/c = 0.60 and 0.80 

at all angles of attack. Modified ESP module 

pressure measurement accuracies of 

approximately +/-1.6 percent would bracket the 

majority of ∆Cp values up to α = 24 degrees. 

Comparison of the pressure coefficients to Cp*

reveals a small pocket of supersonic flow 

underneath the porous LEX and wing vortex 

flows for both repeat runs at α = 24 degrees and 

x/c = 0.60 (figure 28(d)). The data scatter 

between repeat runs is less pronounced at 

Mach = 0.85 (figure 29), and the maximum ∆Cp

values generally do not exceed approximately 

0.8 percent F. S. range of the ESP module. As 

noted with the solid LEX, there is an apparent 

bias in the delta coefficient plots. In this case, 

the ∆Cp values are almost exclusively negative 

at α = 16, 20, and 24 degrees. Comparison of 

the pressure coefficients to the critical value at 

Mach = 0.85 indicates that the wing upper 

surface is dominated by supersonic flow 

conditions.

Force and Moment Measurements

Figure 30 and figure 31 show the delta 

coefficient plots of the normal force, axial force,

and pitching moment coefficients (∆CN, ∆CA, 

∆Cm) versus α corresponding to the solid LEX at 

Mach = 0.50 and 0.85, respectively. The 

95 percent confidence limits for the 755 balance 

accuracies derived from the balance calibration 

expressed in terms of the aerodynamic 

coefficients at Mach = 0.50 and 0.85 are 

depicted as horizontal dashed lines in the delta 

coefficient plots. Computation of the delta 

aerodynamic coefficients was performed at

angles of attack from 2 degrees to 28 degrees in 

2-degree increments. At Mach = 0.50, the 

majority of the delta normal force coefficient 

values are reasonably bounded by the balance 

calibration accuracy limit of 0.05 percent F. S., 

and all delta coefficient values are bracketed by 

a modified limit of approximately 0.1 percent 

F. S. The axial force coefficient scatter is 

outside the balance accuracy limit of 0.26 

percent F. S. at all angles of attack. However, a 

modified limit corresponding to approximately 

0.36 percent F. S. effectively contains all delta 

axial force coefficient values at this Mach 

number. Similarly, the pitching moment 

coefficient scatter typically exceeds the balance 

calibration accuracy limit of 0.09 percent F. S., 

but the delta coefficient values are bracketed by 

a revised limit of approximately 0.16 percent 

F. S. Similar results are obtained at 
17

Mach = 0.85 (figure 31), where modified 

balance calibration accuracy limits of 

0.1 percent F. S., 0.42 percent F. S., and 

0.5 percent F. S. capture the ∆CN, ∆CA, ∆Cm

values, respectively, across the entire range of 

angle of attack from 2 to 28 degrees.

Figure 32 and figure 33 show the delta 

coefficient plots of the normal force, axial force, 

and pitching moment coefficients (∆CN, ∆CA, 

∆Cm) versus α corresponding to the porous LEX 

at Mach = 0.50 and 0.85, respectively. The 

balance calibration accuracy limits for normal 

force and axial force are reasonable estimators 

of the overall delta coefficient scatter at 

Mach = 0.50 (figure 32). The pitching moment 

coefficient scatter exceeds the corresponding 

balance accuracy limit at all angles of attack but 

does not exceed 0.25 percent F. S. across the 

range of angle of attack. Increased data scatter 

is apparent at Mach = 0.85 (figure 33), and the 

delta coefficient values typically exceed the 

balance calibration accuracy limits for the 

normal force, axial force, and pitching moment 

coefficients. However, modified balance 

calibration accuracy limits corresponding to 

approximately 0.21 percent F. S., 0.48 percent

F. S., and 0.54 percent F. S. serve as reasonable

estimates of the overall data scatter for the 

normal force, axial force, and pitching moment 

coefficients in repeat runs at this Mach number. 

Note that the initial run at Mach = 0.85 did not 

obtain data at α = 26 and 28 degrees, so delta 

coefficient values were not determined at these 

angles of attack.

The within-test data repeatability assessment 

typically reveals data scatter that exceeds the 

pressure and balance measurement 

instrumentation accuracies derived from pre-test 

calibrations. However, the data repeatability is 

considered reasonable given the known 

topography of the flow field, which is dominated 

by leading-edge flow separation, vortex flows, 

and shock waves at the transonic speeds. This 

assessment provides a reference with which to 

assess the major configuration and test condition 

changes that are the primary focus of this report. 

Tunnel-to-Tunnel Data Reproducibility

Pressure Distributions

A rigorous examination of tunnel-to-tunnel 

data reproducibility is provided in figure 34

through figure 40, which compare the wing 

upper surface static pressure distributions 

obtained on several configurations in 

8-Foot TPT Test 1057 at Mach = 0.50 and a unit 

Reynolds number of 2.5 million and in 

7- by 10-Foot HST Test 202 at Mach = 0.50 and 

a unit Reynolds number of 3.0 million. Each 

figure contains tunnel-to-tunnel comparisons of 

the pressure distributions at nominal angles of 

attack of 10 degrees and 20 degrees. Seven 

common configurations were tested in these 

facilities corresponding to (1) solid LEX with 

tails off (figure 34), (2) porous LEX (centerline 

tail on in Test 1057, tail off in Test 202) 

(figure 35), (3) wing only (centerline tail on in 

Test 1057, tail off in Test 202) (figure 36), 

(4) solid LEX with twin uncanted tails 

(figure 37), (5) solid LEX with twin canted tails

(figure 38), (6) porous LEX with twin uncanted 

tails (figure 39), and (7) wing only with twin 

canted tails (figure 40). At zero sideslip 

conditions, the presence of the centerline tail did 

not affect the wing pressure distributions at 

x/c = 0.30, 0.60, and 0.80 and α = 10 and 

20 degrees (reference 4). Consequently, 

comparisons of the pressure distributions 

obtained with the porous LEX and wing only 

with centerline tail in Test 1057 to the 

corresponding configurations with centerline tail 

off in Test 202 were considered valid. The 

vortex-dominated surface pressure signatures 

obtained on all configurations at α = 10 and 

20 degrees exhibit reasonable tunnel-to-tunnel 

agreement. However, the scatter in the pressure 

data is generally greater in the tunnel-to-tunnel 

comparisons compared to the within-test data 

repeatability scatter observed at Mach = 0.50 in 

the previous section. This is not an unexpected 

result, since the 8-Foot TPT and 7- by 10-Foot 

HST tunnel tests were conducted over a time 

period of two years and featured different 

pressure and force and moment measurement 

instrumentation, model support systems, and 

testing environments. 
18

The data presented in figure 34 through

figure 40 were used to create the delta pressure 

coefficient plots in figure 41 through figure 47. 

The delta coefficients were computed in these 

plots as ∆Cp = Cp(initial point) - Cp(repeat point)

= Cp(8-Foot TPT) - Cp(7- by 10-Foot HST). The 

ESP module pressure measurement accuracies 

representing the upper and lower 95 percent

confidence limits at Mach = 0.50 corresponding 

to the test condition in the 8-Foot TPT are 

displayed as horizontal dashed lines in the delta 

pressure coefficient plots. In addition, the 

pressure coefficient corresponding to the local 

speed of sound, Cp*, at Mach = 0.50 is depicted 

as a horizontal dashed line in selected plots of 

the upper surface static pressure coefficient

distributions obtained in both facilities. The 

tunnel-to-tunnel delta coefficient values in 

figure 41 through figure 47 are generally well 

outside the limits defined by ESP module 

pressure measurement accuracy. Modified 

pressure measurement accuracy limits 

corresponding to approximately +/-1 percent to 

+/-2 percent of the F. S. range of the ESP 

module would effectively bound the majority of 

tunnel-to-tunnel ∆Cp values, depending on the 

configuration and the angle of attack. This level 

of tunnel-to-tunnel repeatability is considered 

satisfactory given the known topology of the 

flow field. An exception to this is observed in 

figure 43 corresponding to the wing only (with 

centerline tail in Test 1057 and with tail off in 

Test 202), where the difference in peak suction 

pressure coefficients at x/c = 0.30 exceeds 

5 percent of the F. S. range of the ESP module. 

The wing-only configurations presented in 

figure 43 and in figure 47 (with twin canted 

tails) exhibit peak suction pressure coefficients 

that exceed the critical pressure coefficient, Cp*, 

in both facilities. The delta coefficient values in 

figure 41 through figure 47 frequently exhibit a 

systematic variation within a given pressure 

measurement station such that the majority of 

∆Cp values are of the same sign (typically 

negative). This systematic variation could arise 

from many factors, including thermal offsets in 

the ESP measurements, model differences, 

errors in the sting deflection and accelerometer 

calibrations, undocumented misalignments in the 

model, balance, and support system installation, 

differences in support system interference 

effects, and in the tunnel conditions 

(e.g. Reynolds number, humidity level). With 

the exception of the porous LEX configuration, 

the tunnel-to-tunnel delta coefficient values 

are generally greater than the ∆Cp values 

previously presented in the section Within-Test 

Data Repeatability by at least a factor of 2. 

However, the results presented in this section are 

considered a more realistic assessment of the 

data reproducibility on the 65-degree cropped 

delta wing model, since it accounts for 

numerous sources of variation that are not 

manifested in the within-run-series repeats.

Force and Moment Measurements

Figure 48 through figure 54 compare the 

longitudinal aerodynamic coefficients obtained 

on the seven common configurations described 

in the previous section Pressure Distributions. 

The seven configurations correspond to (1) solid 

LEX with tails off (figure 48), (2) porous LEX 

(centerline tail on in Test 1057, tail off in Test 

202) (figure 49), (3) wing only (centerline tail on 

in Test 1057, tail off in Test 202) (figure 50), 

(4) solid LEX with twin uncanted tails 

(figure 51), (5) solid LEX with twin canted tails 

(figure 52), (6) porous LEX with twin uncanted 

tails (figure 53), and (7) wing only with twin 

canted tails (figure 54). The range of angle of 

attack in the 8-Foot TPT corresponded 

approximately to +2 degrees to a maximum of 

28 to 30 degrees depending on the configuration. 

Angle of attack increments were typically 2 

degrees in Test 1057. An expanded range of 

angle of attack was possible using the high 

angle-of-attack support system in the 7- by 10-

Foot HST, although the maximum angle of 

attack was typically limited by model and/or 

balance dynamics to 32 to 40 degrees depending 

on the configuration. Nominal angle of attack 

increments of 2.5 degrees were used in Test 202. 

Figure 48 (solid LEX with tails off) shows 

reasonable tunnel-to-tunnel agreement in the lift, 

drag, and pitching moment coefficients within 

the overlapping ranges of the angle of attack and 

to the scale of the data plots. Test data were 

obtained in the 7- by 10-Foot HST beyond the 
19

angle of attack for onset of leading-edge vortex 

breakdown over the wing (reference 4), which 

promoted the discontinuities in the lift, drag, and 

pitching moment coefficients at an angle of 

attack of approximately 32.5 degrees. Good 

agreement is also observed in figure 49 (porous 

LEX (centerline tail on in Test 1057, tail off in 

Test 202)). Testing in the 7- by 10-Foot HST 

was not conducted at angles of attack greater 

than 32.5 degrees because of dynamic overload 

of the balance axial force component. This 

phenomenon was assumed to be caused by 

unsteady flow through the porous LEX surface. 

The test results obtained on the wing only 

(centerline tail on in Test 1057, tail off in Test 

202) in figure 50 show good tunnel-to-tunnel 

agreement, including the onset of vortex 

breakdown near the wing trailing edge 

beginning at an angle of attack of approximately 

26 degrees (reference 4). The solid LEX with 

twin uncanted tails (figure 51) and twin canted 

tails (figure 52) promote more complex flow 

fields involving interactions of the leading-edge 

vortices and the wing-mounted tail surfaces. 

These interactions cause multiple nonlinearities 

in the lift, drag, and pitching moment 

coefficients which are effectively captured in 

both facilities. Tunnel-to-tunnel differences 

exist in the coefficient values at the higher 

angles of attack. This could be caused by vortex 

flow hysteresis (reference 4) arising from the 

different angle-of-attack schedules in the two 

tunnels. Evidence of potential hysteresis effects 

is also present in the tunnel-to-tunnel 

comparisons of the porous LEX with twin 

uncanted tails in figure 53. The path taken in 

the angle-of-attack sweep in the 7- by 10-Foot 

HST leads to an apparent earlier vortex 

breakdown over the wing which promotes more 

significant tunnel-to-tunnel differences in the 

lift, drag, and pitching moment coefficients at 

the higher angles of attack. In contrast, the wing 

only with twin canted tails (figure 54) exhibits 

very good tunnel-to-tunnel agreement, including 

replication of the nonlinear aerodynamic 

characteristics promoted by vortex-tail 

interactions.

The data presented in figure 48 and figure 51

through figure 54 were used to create the delta 

coefficient plots of the normal force, axial force, 

and pitching moment coefficients (∆CN, ∆CA, 

∆Cm) versus α in figure 55 through figure 59. 

The 95 percent confidence limits for the 755 

balance accuracies derived from the balance 

calibration expressed in terms of the 

aerodynamic coefficients at Mach = 0.50 in 

Test 1057 are depicted as horizontal dashed lines 

in the delta coefficient plots. Computation of 

the delta aerodynamic coefficients was 

performed in 2-degree increments for selected 

ranges of the angle of attack that were truncated 

to avoid extrapolation of the 8-Foot TPT data 

and interpolation in regions of significant 

aerodynamic nonlinearities where the angle of 

attack schedules in the two tunnels were too 

disparate. It is noted that the spline fits to the 

original coefficient data occasionally yielded 

spurious results in regions of aerodynamic 

nonlinearities. Delta aerodynamic coefficients 

were not computed using the data previously 

presented in figure 49 and figure 50 because the 

8-Foot TPT data were obtained with the 

centerline vertical tail, whereas the 7- by 10-

Foot HST data were acquired without the tail.

The delta normal force coefficient values in 

figure 55 corresponding to the solid LEX with 

tails off are bracketed by a modified limit of 

approximately 0.1 percent F. S. compared to the 

balance calibration accuracy limit of 

0.05 percent F. S. The axial force coefficient 

scatter is within the balance accuracy limit of 

0.26 percent F. S. at all angles of attack for 

which tunnel-to-tunnel delta coefficients were 

computed. The pitching moment coefficient 

exhibits the largest data scatter, and the tunnelto-tunnel delta coefficients are bracketed by a 

modified limit of approximately 1 percent F. S. 

compared to the balance calibration accuracy 

limit of 0.09 percent F. S. The remaining four 

configurations in figure 56 through figure 59

feature the twin uncanted or canted tails, and the 

tunnel-to-tunnel data scatter is sensitive to the 

interactions of the vortex flows with the tail 

surfaces. For example, the delta normal force 

coefficient scatter for the solid LEX with twin 
20

uncanted tails (figure 56) is bracketed by a 

modified limit of 0.15 percent F. S. (compared 

to the balance calibration accuracy limit of 

0.05 percent F. S.) except at the highest angles 

of attack where vortex breakdown is assumed to 

occur. At these high angle-of-attack conditions, 

the data scatter is bounded by a modified limit of 

approximately 0.45 percent F. S. The axial force 

coefficient scatter is contained within the 

balance calibration accuracy limit of 

0.26 percent F. S. The delta pitching moment 

coefficient values generally increase slightly as 

the angle of attack increases, and the data scatter 

is effectively contained by a modified limit of 

0.6 percent F. S. The delta normal force 

coefficient values for the solid LEX with twin 

canted tails (figure 57) exhibit a systematic 

variation with the angle of attack, which could 

be caused by slight tunnel-to-tunnel differences 

in the interactions of the vortex flows with the 

canted tails at moderate and high angles of 

attack. The normal force coefficient data scatter 

is contained by a modified limit of 

0.4 percent F. S. compared to the balance 

calibration accuracy limit of 0.05 percent F. S. 

Axial force coefficient data scatter is generally 

within the balance calibration accuracy limit of 

0.26 percent F. S., although a modified limit of 

0.33 percent F. S. brackets all of the delta 

coefficient values for which tunnel-to-tunnel 

comparisons are made. This configuration also 

exhibited the largest delta pitching moment 

coefficient values of the five configurations 

shown in figure 55 through figure 59. An 

expanded limit of 1.8 percent F. S. contains the 

tunnel-to-tunnel data scatter, which is 20 times 

greater than the balance calibration accuracy 

limit of 0.09 percent F. S. Of the five 

configuration comparisons, the porous LEX with 

twin uncanted tails (figure 58) exhibits the 

largest overall data scatter at the higher angles of 

attack. Most of the delta normal force and axial 

force coefficient values are within the respective 

balance calibration accuracy limits except at 

angles of attack of 24 degrees and higher where 

vortex breakdown is assumed to occur. 

Modified limits of approximately 0.8 percent 

F. S. and 0.7 percent F. S. bracket all delta 

normal force and axial force coefficient values, 

respectively, at all angles of attack for which 

tunnel-to-tunnel comparisons are made. 

Similarly, the delta pitching moment coefficients 

are bracketed by a modified limit of 0.8 percent

F. S. except at the highest angle of attack where 

the tunnel-to-tunnel difference nearly doubles. 

The final configuration comparison is shown in 

figure 59, which corresponds to the wing alone 

with the twin canted tails. The normal force 

coefficient data scatter is within 0.1 percent F. S. 

except at an angle of attack of 18 degrees, where 

subtle differences in the vortex-tail interaction in 

the two tunnels is assumed to promote a 

two-fold increase in the data scatter. The delta 

axial force coefficient values are bracketed by 

the balance calibration accuracy limit. A 

gradual increase in the pitching moment 

coefficient data scatter is observed as the angle 

of attack increases. The majority of delta 

coefficient values are bounded by a modified 

limit of 0.2 percent F. S., although a limit of 

0.4 percent F. S. is required to bracket the data 

scatter at the higher angles of attack.

Basic Configuration Aerodynamics

The pressure distributions and six-component 

force and moment coefficients obtained on the 

four primary configurations in 8-Foot TPT 

Test 1057 corresponding to the wing-alone, solid 

LEX-wing, porous LEX-wing, and canard-wing 

are presented in this section. All configurations 

were tested with the centerline vertical tail. Test 

results are presented for each configuration in 

the format of Mach number sweeps, angle-ofattack sweeps, and sideslip sweeps. The data 

presented in this section are intended to establish 

the basic aerodynamic characteristics of each 

configuration to facilitate the discussion of the

configuration comparisons in later sections of 

this report.

Pressure Distributions

Wing Alone

Mach Number Sweeps

Figure 60 presents the effect of the Mach 

number on the upper surface static pressure 

distributions obtained on the wing-alone
21

configuration at selected angles of attack of 12, 

16, and 20 degrees. Test results were obtained 

at nominal Mach numbers of 0.50, 0.85, and 

1.20. At each angle of attack, increasing the 

Mach number decreases the maximum vortexinduced suction pressure coefficient at a given 

measurement station, flattens the pressure 

distributions, and promotes an inboard shift of 

the suction pressure peak. These results are 

consistent with data obtained in previous 

experiments of slender wings and, specifically, 

with the 65-degree cropped delta model

(reference 4).

Angle-of-Attack Sweeps

Figure 61 through figure 63 show the 

pressure distributions obtained in angle-of-attack 

sweeps at Mach = 0.50, 0.85, and 1.20, 

respectively. The pressure distributions at 

Mach = 0.50 in figure 61 are characterized by a 

single, pronounced suction pressure peak at each 

measurement station that increases in magnitude 

and moves inboard as the angle of attack is 

increased up to the onset of vortex breakdown 

over the wing. Small regions of locally 

supersonic flow exist under the vortex at 

x/c = 0.30 and angles of attack of 20 degrees and 

higher and at x/c = 0.60 and angles of attack of 

24 degrees and higher, since the suction pressure 

coefficients exceed the critical value, 

Cp

* = -2.133 at Mach = 0.50. The upstream 

influence of vortex breakdown over the wing is 

detected in the pressure distributions at 

x/c = 0.60 and 0.80 at α = 28 degrees, since the 

peak suction pressure coefficients exhibit a 

decline compare to the results obtained at 

α = 24 degrees. Although vortex breakdown is 

inferred upstream of the wing trailing edge, 

there is no indication that it has progressed to the 

aft pressure measurement at α = 28 degrees.

Compared to the results shown in figure 61 at 

Mach = 0.50, the suction pressure peaks are less 

pronounced and the suction peak magnitude 

exhibits less variation from the forward to the aft 

pressure measurement stations at Mach = 0.85 in 

figure 62. Large regions of locally supersonic 

flow exist at all three measurement stations, 

since the pressure coefficients exceed the critical 

value of Cp

* = -0.302 at Mach = 0.85. Evidence 

of a rapid forward advancement of vortex 

breakdown is apparent in the pressure 

distributions at x/c = 0.60 and 0.80 at angles of 

attack greater than 20 degrees. The marked 

decline in the maximum suction pressure 

coefficients and the flattening of the pressure 

distributions at α = 24 and 28 degrees are 

consistent with the passage of vortex breakdown 

at these measurement stations. Reference 3 and 

reference 4 have shown that the development of 

a strong normal shock wave over the wing 

causes early vortex breakdown and a more rapid

forward progression of the breakdown location 

at Mach = 0.85.

The pressure coefficients exhibit a more 

conical and broader distribution (reference 4) at 

Mach = 1.20 in figure 63. Vortex breakdown 

does not occur up to the highest angle of attack 

of 22 degrees (constrained by balance normal 

force design limits). At α = 22 degrees, the peak 

suction pressure coefficients approach 90 

percent of the vacuum pressure coefficient, 

Cp,v = -0.992, at Mach = 1.20.

Sideslip Sweeps

Figure 64 and figure 65 illustrate the pressure 

distributions obtained in sideslip sweeps at 

α = 12, 16, 20, and 24 degrees at Mach = 0.50 

and 0.85, respectively. Sideslip sweeps were 

conducted at α = 12 and 16 degrees at 

Mach = 1.20 as shown in figure 66. The data 

are plotted at sideslip angles from approximately 

-8 degrees (“nose right”) to +8 degrees (“nose 

left”) in nominal 4-degree increments. This 

provided data on the leeward and windward 

wings, since pressure taps were available on 

only the right wing upper surface.

The data obtained at Mach = 0.50 and 

α = 12 degrees in figure 64(a) indicate that 

increasing the sideslip angle from approximately 

-8 degrees to +8 degrees promotes an inboard 

displacement and increased magnitude of the 

peak suction pressure coefficient at a given 

pressure measurement station. This is consistent 

with an inboard and downward movement of the 

wing leading-edge vortex as the sideslip angle 

increases from negative to positive values. In 

general, the windward wing exhibits higher peak 
22

suction pressure levels compared to the leeward 

wing at x/c = 0.30, 0.60, and 0.80. This is not 

the case at α = 16 degrees, however, because of 

the onset of vortex breakdown over the 

windward wing. This is manifested in the 

pressure distributions at x/c = 0.80 in 

figure 64(b) by a reduced suction pressure peak 

and a flattening of the overall pressure 

distribution at a maximum positive sideslip 

angle of β = +7.66 degrees at this angle of 

attack. The effect of vortex breakdown on the 

windward wing pressure distributions advances 

forward to x/c = 0.60 at α = 20 and 24 degrees 

in figure 64(c) and figure 64(d), respectively. 

Although the maximum suction pressure levels 

on the windward wing are typically less than the 

leeward wing, the burst vortex tends to induce 

higher suction pressure levels along the inboard 

region of the windward wing.

The pressure distributions at Mach = 0.85 in 

figure 65 exhibit trends that are similar to those 

observed at Mach = 0.50 in figure 64. However, 

the forward advance of vortex breakdown over 

the windward wing is more dramatic at this low 

transonic Mach number because of the 

development of a strong normal shock wave. 

The pressure distributions on the leeward wing 

do not exhibit this effect, presumably because of 

the higher effective wing sweep and a mitigation 

of shock-induced vortex breakdown 

(reference 4). 

Sideslip sweeps were constrained to a 

maximum angle of attack of 16 degrees at 

Mach = 1.20 because of balance design load 

limits. Compared to the data obtained at 

Mach = 0.50 and 0.85, the pressure distributions

at Mach = 1.20 in figure 66 exhibit lower overall 

suction pressure levels and flatter distributions,

which are consistent with a weaker and more 

elliptically-shaped leading-edge vortex at this 

higher transonic Mach number (reference 4). 

Vortex breakdown did not occur at any 

combination of angle of attack and sideslip that 

was tested at this Mach number in Test 1057 

(see also reference 4). The effect of increasing 

the sideslip angle from negative to positive 

values at Mach = 1.20 is to promote a broader 

“rooftop” pressure distribution and marginally 

higher maximum suction pressure coefficients 

on the windward wing compared to the leeward 

wing. 

Solid LEX

Mach Number Sweeps

Figure 67 presents the effect of the Mach 

number on the upper surface static pressure 

distributions obtained with the solid LEX at 

selected angles of attack of 12, 16, 20, and 24

degrees. Test results were obtained at Mach 

numbers of 0.50, 0.85, and 1.20. At α = 12 and 

16 degrees (figure 67(a) and figure 67(b)), the 

pressure distributions at x/c = 0.60 and 0.80 

reveal the signatures of the LEX and wing 

vortices, which do not directly interact at this 

angle of attack (reference 4). The wing vortex 

induces the more prominent suction pressure 

peak, whereas a more subtle pressure signature 

situated farther inboard on the wing is induced 

by the LEX vortex. The suction pressure 

coefficients are subcritical at all measurement 

stations up to α = 24 degrees. The character of 

the pressure distributions does not change 

significantly as the Mach number increases from 

0.50 to 0.85. However, increasing the Mach 

number to 1.20 typically decreases the overall 

vortex-induced suction pressure levels, flattens 

the pressure distributions, and promotes an 

inboard shift of the wing vortex suction pressure 

peak in a manner similar to that observed on the 

wing-alone. This is particularly apparent at 

α = 20 and 24 degrees in figure 67(c) and 

figure 67(d). The pressure distributions at 

α = 20 and 24 degrees are characterized by a 

single suction pressure peak as the LEX and 

wing vortices directly interact. In addition, 

vortex breakdown does not occur up to the 

highest angle of attack shown in figure 67. 

These results are also consistent with data 

obtained in previous testing of the solid LEXwing configuration (reference 4).

Angle-of-Attack Sweeps

The direct interaction of the vortices is 

characterized by an intertwining of the LEX and 

wing vortex cores at the higher angles of attack. 

This interaction and the absence of vortex 
23

breakdown can be inferred from the angle of 

attack sweep at Mach = 0.50 in figure 68. The 

values of the pressure coefficients are generally 

subcritical, except at α = 28 degrees and 

x/c = 0.80 where the suction pressure peak is 

slightly greater than the critical value, 

Cp

* = -2.133, at Mach = 0.50. The trends are 

similar in angle of attack sweeps at Mach = 0.85

and 1.20 in figure 69 and figure 70, respectively,

although the suction pressure distributions are 

flatter and the rate of increase of the suction 

pressure peaks is constrained. Vortex 

breakdown is not manifested in the pressure 

distributions at Mach = 0.85 and 1.20. The

pressure coefficients are generally supercritical 

at the higher angles of attack at Mach = 0.85 

(Cp

* = -0.302). At Mach = 1.20, the maximum 

suction pressure coefficient at x/c = 0.80 is 

approximately 80 percent of the vacuum 

pressure coefficient at this Mach number 

(Cp,v = -0.992).

Sideslip Sweeps

Figure 71 and figure 72 illustrate the pressure 

distributions obtained in sideslip sweeps at 

α = 12, 16, 20, and 24 degrees at Mach = 0.50 

and 0.85, respectively. Sideslip sweeps were 

conducted at α = 12, 16, and 20 degrees at 

Mach = 1.20 as shown in figure 73. The data 

obtained at Mach = 0.50 indicate that increasing 

the sideslip angle from approximately -8 degrees 

to +8 degrees promotes an inboard displacement 

and increased magnitude of the peak suction 

pressure coefficients at a given pressure 

measurement station at all angles of attack. In 

general, direct interaction of the LEX and wing 

vortices is more pronounced on the leeward 

wing, corresponding to negative sideslip angles, 

which yields a single suction peak in the 

pressure distributions. Conversely, the LEX and 

wing vortex interaction is delayed to higher 

angles of attack on the windward wing, 

corresponding to positive sideslip angles, with a 

concurrent persistence of dual suction peaks in 

the pressure distributions. There is no indication 

of vortex breakdown in the pressure 

distributions up to the highest angle of attack of 

24 degrees in figure 71. Similar results are 

observed at Mach = 0.85 and 1.20 in figure 72

and figure 73, respectively, although the effects 

are more subtle because of the broader pressure 

distributions and reduced suction pressure peaks 

at the higher Mach numbers.

Porous LEX

Mach Number Sweeps

Figure 74 presents the effect of the Mach 

number on the upper surface static pressure 

distributions obtained with the porous LEX at 

selected angles of attack of 12, 16, 20, and 24 

degrees. Test results were obtained at Mach 

numbers of 0.50, 0.85, and 1.20. The porous 

LEX configuration is characterized by a single 

vortex-induced pressure signature at all Mach 

numbers and angles of attack shown in 

figure 74. Similar to the results obtained on the 

wing-alone configuration in figure 60, the 

suction pressure peak at each measurement 

station moves inboard and the pressure 

distributions are flatter as the Mach number 

increases. The peak suction pressure coefficient

at x/c = 0.60 exceeds the critical pressure 

coefficient, Cp

* = -2.133, at Mach = 0.50 and 

α = 24 degrees. At Mach = 0.85, the surface 

pressure field is dominated by supersonic flow, 

since the pressure coefficients are typically 

greater than the critical value, Cp

* = -0.302. 

Angle-of-Attack Sweeps

Figure 75 through figure 77 present the upper 

surface static pressure coefficient distributions in 

angle-of-attack sweeps at Mach = 0.50, 0.85, 

and 1.20, respectively. The pressure 

distributions at Mach = 0.50 in figure 75

typically exhibit a single suction peak at each 

measurement station. The suction peaks 

increase in magnitude and the character of the 

pressure distributions is unchanged as the angle 

of attack increases to the maximum value of 

α = 28 degrees in figure 75. These results 

suggest vortex breakdown has not occurred over 

the wing. Small pockets of locally supersonic 

flow exist under the vortex system on the porous 

LEX-wing at α = 24 and 28 degrees and 

x/c = 0.60 and 0.80. The data at Mach = 0.85 

and 1.20 in figure 76 and figure 77, respectively, 

reveal broader, single-peaked pressure 

distributions and overall suction pressure levels 

that consistently increase with increasing angle 
24

of attack. Similar to the results obtained on the 

solid LEX-wing configuration in figure 69 and 

figure 70, vortex breakdown is not manifested in 

the pressure distributions at Mach = 0.85 and 

1.20. The pressure coefficients are generally 

supercritical at the higher angles of attack at 

Mach = 0.85 (Cp

* = -0.302). At Mach = 1.20, 

the maximum suction pressure coefficient at 

x/c = 0.80 is approximately 80 percent of the 

vacuum pressure coefficient at this Mach 

number (Cp,v = -0.992).

Sideslip Sweeps

Figure 78 and figure 79 illustrate the pressure 

distributions obtained in sideslip sweeps at 

α = 12, 16, 20, and 24 degrees at Mach = 0.50 

and 0.85, respectively. Sideslip sweeps were 

conducted at α = 12, 16, and 20 degrees at 

Mach = 1.20 as shown in figure 80. The data

trends at Mach = 0.50 in figure 78 are similar to 

those observed on the solid-LEX in figure 71. 

Specifically, increasing the sideslip angle from 

approximately -8 degrees to +8 degrees typically

promotes an inboard displacement and increased 

magnitude of the peak suction pressure 

coefficients at a given pressure measurement 

station at all angles of attack. In addition, there 

is no indication of vortex breakdown up to the 

highest angle of attack of 24 degrees in 

figure 78. In contrast to the solid LEX-wing 

results, distinct LEX and wing vortex pressure 

signatures are not discernible in the pressure 

distributions. Similar trends are observed at 

Mach = 0.85, α = 12 and 16 degrees 

(figure 79(a) and figure 79(b)), and x/c = 0.30 

and 0.60. At x/c = 0.80, increasing the sideslip 

angle from approximately -8 degrees to 

+8 degrees does not yield a consistent increase 

in the peak suction pressures. At α = 24 degrees 

(figure 79(d)), the windward pressure 

distributions at the β = +7.20 degrees and 

x/c = 0.60 and 0.80 exhibit a marked decline in 

the overall suction pressure levels compared to 

the corresponding pressure distributions at 

β = +4.01 degrees. These results suggest the 

onset of shock-induced vortex breakdown over 

the windward wing in a manner similar to the 

wing-alone configuration in figure 65. The 

pressure distributions at Mach = 1.20 in 

figure 80 are less sensitive to the sideslip angle 

at α = 12, 16, and 20 degrees. In general, the 

overall suction pressure level is higher on the 

windward wing compared to the leeward wing. 

Vortex breakdown does not occur at this Mach 

number up to the highest angle of attack in 

figure 80. 

Canard

Mach Number Sweeps

Figure 81 presents the effect of the Mach 

number on the upper surface static pressure 

distributions obtained with the canard at selected 

angles of attack of 12, 16, 20, and 24 degrees. 

The canard generates a leading-edge vortex that 

favorably interacts with the wing flow field 

(reference 3 and reference 14). Unlike the solid 

LEX-wing configuration, however, the direct 

interaction, or intertwining, of the canard and 

wing vortices is mitigated, since the canard 

vortex is situated higher above the wing surface 

compared to the LEX vortex. Consequently, the 

signature of the canard vortex is not directly 

observed in the pressure distributions, which are 

characterized by a single suction pressure peak 

induced by the wing leading-edge vortex at all 

Mach numbers and angles of attack shown in 

figure 81. The pressure distributions at 

x/c = 0.30 show a consistent decrease in the 

peak suction pressure coefficients at all angles of 

attack as the Mach number increases from 0.50 

to 1.20. This is not the case at x/c = 0.60 and 

0.80 and α = 12 and 16 degrees (figure 81(a)

and figure 81(b), respectively), however, where 

the decline in the peak suction pressures does 

not occur until Mach = 1.20. The peak suction 

pressure coefficients exhibit a more consistent 

decrease with increasing Mach number at these 

measurement stations and α = 20 and 24 degrees 

(figure 81(c) and figure 81(d), respectively). 

However, the overall suction pressure levels and 

the location of the suction pressure peaks are 

dependent on the Mach number. For example, 

the pressure distributions at Mach = 0.85 and 

x/c = 0.60 and 0.80 generally exhibit higher 

overall suction pressure levels and an inboard 

shift in the peak suction pressure coefficient 

compared to Mach = 0.50. At Mach = 1.20, 

however, the overall suction pressure levels are 

markedly lower, but the location of the peak 
25

suction pressure level is either unchanged or 

displaced slightly outboard compared to 

Mach = 0.85. These results may reflect 

differences in the mutual interaction of the 

canard and wing vortices depending on the angle 

of attack and Mach number. The peak suction 

pressure coefficient is subcritical at all angles of 

attack at Mach = 0.50, whereas the pressure 

coefficients at Mach = 0.85 are typically greater 

than the critical value, Cp

* = -0.302. 

Angle-of-Attack Sweeps

Figure 82 through figure 84 present the upper 

surface static pressure coefficient distributions in 

angle-of-attack sweeps at Mach = 0.50, 0.85, 

and 1.20, respectively. The pressure 

distributions at Mach = 0.50 in figure 82

typically exhibit a single suction peak at each 

measurement station. The suction peaks 

increase in magnitude as the angle of attack 

increases to the maximum value of α = 28 

degrees in figure 82. The values of the pressure 

coefficients are subcritical at all angles of attack. 

The character of the pressure distributions is 

unchanged up to α = 20 degrees. The pressure 

distributions broaden at α = 24 and 28 degrees, 

which may be caused by a stronger interaction 

of the canard and wing vortices and a resultant 

migration of the wing vortex away from the 

surface. Vortex breakdown effects are not 

observed at any angle of attack in figure 82. 

These results suggest vortex breakdown has not 

occurred over the wing. Similar results are 

obtained at Mach = 0.85 in figure 83, although 

the broadening of the pressure distributions 

occurs at a lower angle of attack 

(α = 20 degrees). The values of the pressure 

coefficients are primarily supercritical 

(Cp

* = -0.302) at all measurement stations at this 

Mach number. There is no indication of 

shock-induced vortex breakdown up to the 

highest angle of attack of 28 degrees. The data 

at Mach = 1.20 in figure 84 reveal broader 

distributions and overall suction pressure levels 

that consistently increase with increasing angle 

of attack. The suction pressure levels under the 

vortex system approach a plateau at each

measurement station, and the peak suction 

pressure coefficient at x/c = 0.80 is 

approximately 75 percent of the vacuum 

pressure coefficient at this Mach number 

(Cp,v = -0.992).

Sideslip Sweeps

Figure 85 and figure 86 illustrate the pressure 

distributions obtained in sideslip sweeps at 

α = 12, 16, 20, and 24 degrees and Mach = 0.50 

and 0.85, respectively. Sideslip sweeps were 

conducted at α = 12, 16, and 20 degrees at 

Mach = 1.20 as shown in figure 87. The data 

trends at Mach = 0.50 in figure 85 typically 

show that increasing the sideslip angle from 

approximately -8 degrees to +8 degrees

promotes an inboard displacement and increased 

magnitude of the peak suction pressure 

coefficients, and overall higher suction pressure 

levels at a given pressure measurement station at 

all angles of attack. There is no indication of 

vortex breakdown up to the highest angle of 

attack of 24 degrees in figure 85. At 

Mach = 0.85 (figure 86), increasing the sideslip 

angle from approximately -8 degrees to 

+8 degrees generally promotes an overall 

increase in the suction pressure levels, although 

the magnitude of the peak suction pressure 

coefficients is constrained. In addition, the 

decline in the peak suction pressure coefficient 

and overall suction pressure level as the sideslip 

angle increases from +4.0 degrees to the 

maximum positive value (+7.29 degrees) 

suggests the presence of shock-induced vortex 

breakdown effects at α = 24 degrees and 

x/c = 0.80 (figure 86(d)). The pressure 

distributions at Mach = 1.20 in figure 87 are less 

sensitive to the sideslip angle at α = 12, 16, and 

20 degrees. In general, the overall suction 

pressure level is higher on the windward wing

compared to the leeward wing. Vortex 

breakdown does not occur at this Mach number 

up to the highest angle of attack in figure 87.

Force and Moment Measurements

Wing Alone

Mach Number Effects

Figure 88 presents the effect of the Mach 

number on the longitudinal aerodynamic 

characteristics of the wing-alone configuration 

with centerline vertical tail. The local lift curve 
26

slope at Mach = 0.50 decreases at angles of 

attack between 26 and 28 degrees because of the 

onset of vortex breakdown upstream of the wing 

trailing edge. The latter was inferred from the 

wing-alone pressure distributions in angle of 

attack sweeps in figure 61. The lift curve slope 

is higher at Mach = 0.85 up to an angle of attack 

of approximately 23 degrees. The pressure 

distributions in figure 60 suggested an overall 

increase in the suction pressure levels at 

Mach = 0.85 compared to Mach = 0.50. The 

breaks in the lift, drag, and pitching moment 

coefficient curves at an angle of attack between 

23 and 24 degrees are caused by the onset of 

shock-induced vortex breakdown over the wing. 

Increasing the Mach number to 1.20 decreases 

the lift curve slope, increases the drag, and 

promotes a stable shift in the pitching moment 

curve (nose-down increments). The Mach 

number effect on the pitching moment curve is a 

result of an aft shift in the center of pressure 

(reference 4).

Sideslip Effects

Figure 89 through figure 91 present the 

lateral-directional aerodynamic characteristics at

Mach = 0.50, 0.85, and 1.20, respectively, in 

sideslip sweeps at selected angles of attack. The 

rolling moment coefficient is of particular 

interest in these plots. The wing-alone 

configuration exhibits a stable variation of the 

rolling moment coefficient with the sideslip 

angle at Mach = 0.50 (figure 89) and angles of 

attack of 12 and 16 degrees. Unstable breaks 

occur in the rolling moment curves at α = 16 and

20 degrees between β = +3 to +4 degrees and 

β = -3 to -4 degrees, which are caused by the 

onset of vortex breakdown over the windward 

wing. This effect was previously inferred from 

the pressure distributions shown in figure 64. 

This effect is apparent at smaller sideslip angles 

at α = 24 degrees, where the unstable breaks 

occur between β = +1 to +2 degrees and 

β = -1 to -2 degrees. Corresponding 

nonlinearities occur in the yawing moment 

coefficient variation with sideslip angle, which 

exhibits reduced directional stability at α = 20 

and 24 degrees caused by the passage of vortex 

breakdown over the windward wind and reduced 

dynamic pressure at the centerline vertical tail.

Shock-induced vortex breakdown on the 

windward wing causes more pronounced 

unstable breaks in the rolling moment curves at 

Mach = 0.85 (figure 90). The more dramatic 

onset of vortex breakdown over the windward 

wing was also indicated in the pressure 

distributions at Mach = 0.85 previously shown 

in figure 65. The rolling moment breaks are 

most apparent at α = 20 degrees between 

β = +3 to +4 degrees and β = -4 to -5 degrees

and at α = 24 degrees between 

β = +2 to +3 degrees and β = -2 to -3 degrees. 

The different onset angles at α = 20 degrees are 

attributed to asymmetries in the vortex-shock 

interactions at the positive and negative sideslip 

angles. Nonlinearities also occur in the yawing 

moment coefficient variation with sideslip angle,

particularly at α = 24 degrees, as a result of an 

interaction of the burst vortex system with the 

centerline vertical tail.

Stable variations of the rolling moment and 

yawing moment coefficients with sideslip are 

apparent at Mach = 1.20 and α = 12 and 16 

degrees in figure 91. These trends are consistent 

with the pressure distributions previously shown 

in figure 66. 

Solid LEX

Mach Number Effects

Figure 92 presents the effect of the Mach 

number on the longitudinal aerodynamic 

characteristics of the solid LEX-wing

configuration with centerline vertical tail. 

Increasing the Mach number from 0.50 to 0.85

promotes a slight increase in the lift coefficient 

at angles of attack from approximately 2 degrees 

to 16 degrees, which is attributed to a slight 

increase in the overall suction pressure levels 

along the rear portion of the wing at the higher 

Mach number (see, for example, figure 67(a)

and figure 67(b)). At angles of attack greater 

than about 16 degrees, increasing the Mach 

number generally decreases the lift coefficient 

and lift curve slope, increases the drag, and 

promotes a stable (nose-down) shift in the 

pitching moment curve. Vortex breakdown does 

not occur at any condition tested at Mach = 0.50 

through 1.20, therefore, there are no 
27

discontinuities in the longitudinal aerodynamic 

characteristics. 

Sideslip Effects 

Figure 93 through figure 95 present the 

lateral-directional aerodynamic characteristics at 

Mach = 0.50, 0.85, and 1.20, respectively, in 

sideslip sweeps at selected angles of attack. 

The solid LEX-wing configuration exhibits a 

stable variation of the rolling moment 

coefficient with the sideslip angle at 

Mach = 0.50 and all angles of attack shown in 

figure 93. In addition, the stable variation of the 

yawing moment coefficient is unaffected by the 

angle of attack. There are no aerodynamic 

discontinuities in the lateral-directional 

aerodynamic characteristics at Mach = 0.50 

because of the absence of vortex breakdown 

effects. The aerodynamic force and moment 

characteristics are consistent with the solid 

LEX-wing pressure distributions shown 

previously in figure 71. Stable variations of the 

rolling moment coefficient with the sideslip 

angle are also observed at Mach = 0.85 and 1.20 

in figure 94 and figure 95, respectively. 

However, reduced levels of directional stability 

occur at Mach = 0.85 and 1.20 at the higher 

angles of attack, which could be associated with 

reduced dynamic pressure and/or adverse 

vortex-induced sidewash at the vertical tail at the 

higher Mach numbers. Evidence of vortex 

breakdown was not observed previously in the 

pressure distributions in figure 73 and figure 74 

at Mach = 0.85 and 1.20, respectively. 

Porous LEX 

Mach Number Effects 

Figure 96 presents the effect of the Mach 

number on the longitudinal aerodynamic 

characteristics of the porous LEX-wing 

configuration with centerline vertical tail. The 

Mach number effects are similar to those 

observed on the solid LEX-wing configuration 

in figure 92. Consequently, no further 

discussion is presented here. 

Sideslip Effects 

Figure 97 through figure 99 present the 

lateral-directional aerodynamic characteristics at 

Mach = 0.50, 0.85, and 1.20, respectively, in 

sideslip sweeps at selected angles of attack. The 

porous LEX-wing configuration exhibits a stable 

variation of the rolling moment coefficient with 

the sideslip angle at Mach = 0.50 and all angles 

of attack shown in figure 97. In addition, the 

stable variation of the yawing moment 

coefficient is relatively unaffected by the angle 

of attack. There are no aerodynamic 

discontinuities in the lateral-directional 

aerodynamic characteristics at Mach = 0.50 

because of the absence of vortex breakdown 

effects. Stable variations of the rolling moment 

coefficient with the sideslip angle are observed 

at Mach = 0.85 in figure 98 except at 

 = 24 degrees and sideslip angles greater than 

+/-6 degrees. The pressure distributions at 

 = 24 degrees shown previously in figure 79(d) 

suggested the onset of shock-induced vortex 

breakdown, which is consistent with the abrupt 

unstable breaks in the rolling moment curve at 

this angle of attack. In addition, the level of 

directional stability decreases at Mach = 0.85 

and  = 20 and 24 degrees. The porous 

LEX-wing configuration exhibits stable 

variations of the rolling moment coefficient with 

sideslip at Mach = 1.20 in figure 99, which is 

consistent with the absence of vortex breakdown 

effects in the pressure distributions in figure 80. 

It is speculated that reduced dynamic pressure 

and/or adverse vortex-induced sidewash at the 

vertical tail is the source of the unstable 

variation of the yawing moment coefficient with 

sideslip at Mach = 1.20 and  = 20 degrees in 

figure 99. 

Canard 

Mach Number Effects 

Figure 100 presents the effect of the Mach 

number on the longitudinal aerodynamic 

characteristics of the canard-wing configuration 

with centerline vertical tail. The canard-wing 

configuration lift coefficient is less sensitive to 

the Mach number compared to the solid 

LEX-wing configuration shown previously in 

figure 92. It is speculated that the absence of a 
28

strong direct interaction of the canard and wing 

vortices renders this configuration less sensitive 

to changes in the Mach number. The drag and 

pitching moment coefficients trends are similar 

to those observed on the solid LEX-wing 

configuration, except for the stable break in the 

pitching moment curve that occurs at a lift 

coefficient of approximately 0.80 corresponding 

to an angle of attack of 16 degrees. This stable 

break is consistent with the pressure 

distributions shown previously in figure 83, 

which showed an overall increase in the suction 

pressure levels at the mid and aft measurement 

stations at Mach = 0.85 as the angle of attack 

increased from 16 to 20 degrees.

Sideslip Effects

Figure 101 through figure 103 present the 

lateral-directional aerodynamic characteristics at 

Mach = 0.50, 0.85, and 1.20, respectively, in 

sideslip sweeps at selected angles of attack. The 

canard-wing configuration exhibits a stable 

variation of the rolling moment coefficient with 

the sideslip angle at Mach = 0.50, 0.85, and 1.20

and at all angles of attack shown in figure 101

through figure 103. The stable variation of the 

rolling moment coefficient through the range of 

sideslip angle at Mach = 0.85 in figure 102 is 

inconsistent with the speculated onset of shockinduced vortex breakdown that was observed in 

the pressure distributions in figure 86(d) at 

α = 24 degrees. It is noted, however, that the 

unstable break and nonlinearity in the yawing 

moment coefficient as the sideslip angle 

increases beyond +/-4 degrees at Mach = 0.85 

and α = 24 degrees in figure 102 is consistent 

with the canard-wing pressure distributions in 

figure 86(d).

Effect of Full LEX Porosity

Laser Vapor Screen Flow Visualization

Mach = 0.85

Figure 104 presents laser vapor screen 

images with the solid LEX, porous LEX, and 

wing alone at Mach=0.85, x/c = 0.80, and angles 

of attack of 16, 20, 24, and 28 degrees. Title 

generator output embedded in each image 

typically contains information related to the test 

conditions; run, polar, and/or configuration 

number; and model attitude. The model 

designation in Test 1057 was “Porous LEX 

Model”, and the name “POR LEX” was an input 

to the title generator setup file. Consequently, 

title generator output containing this information

strictly refers to the model name used in Test 

1057, and the correct configuration designation 

is shown in the individual figure titles.

The vapor screen images of the solid LEX 

configuration are characterized by distinct LEX 

and wing vortical flows having approximately 

circular cross sections that exhibit a strong 

mutual, or direct, interaction, including a coiling 

of the vortices at the higher angles of attack. 

The latter effect causes the vortices to lift away 

from the wing upper surface, which reduces the 

suction pressures induced by the vortical flows 

(reference 4). LEX porosity shifts the 

dominance from the LEX vortex to the wing 

vortex. The LEX vortex is not apparent in any 

of the porous LEX images, which feature a 

single diffused, but stable, wing vortex. 

Porosity does not suppress the LEX vortical 

flow (reference 6); however, it is so weak that 

condensation does not occur in sufficient 

quantity to render it visible in the vapor screen 

images. The porous LEX and wing-alone 

cross-flow patterns at a given angle of attack are 

similar. The most noteworthy difference, 

however, is that the wing vortex with LEX off is 

unstable at the higher angles of attack. For 

example, shock-induced asymmetric vortex 

breakdown occurs at α = 24 degrees, whereas 

the breakdown is symmetric at α = 28 degrees. 

The diffused nature of the wing vortex with the 

porous LEX resembles a burst vortical flow, but 

the video tape recordings of the laser light sheet 

scans along the length of the wing and into the 

near wake indicate an unburst leading-edge 

vortex system. Similar flow visualization results 

were obtained in wind tunnel and flight testing 

of the F/A-18 fighter configuration, which 

exhibited a diffused LEX vortex system caused 

by a LEX upper surface fence for high angle-ofattack vortex control (reference 15).
29

With the exception of shock-induced vortex 

breakdown, the character of the vortex flows on 

the solid LEX, porous LEX, and wing alone at 

Mach = 0.50 are similar to the observations at 

Mach = 0.85 (reference 4). The effects of LEX 

porosity on the cross-flow patterns at 

Mach = 1.20 are also expected to be similar to 

those observed on the same model at lower 

supersonic Mach numbers in reference 6. 

Specifically, LEX porosity also decouples the 

vortex flows at Mach = 1.20. The wing vortex 

cross sections on the solid LEX, porous LEX, 

and wing alone are flatter and more elliptical at 

Mach = 1.20, and vortex breakdown effects are 

assumed to be absent at this Mach number up to 

the maximum angle of attack in Test 1057 

(reference 6).

Pressure Distributions

Mach = 0.50

Figure 105 compares the upper surface static 

pressure coefficient distributions with the solid 

LEX and the LEX with full porosity at 

Mach = 0.50 and angles of attack from 8 degrees 

to 28 degrees in 2-degree increments. The 

wing-alone pressure distributions are included in 

each figure for reference. All configurations 

featured a centerline vertical tail, and all plots 

correspond to a nominal zero-sideslip condition. 

The data corresponding to the solid LEX at 

Mach = 0.50 exhibit dual vortex pressure 

signatures at x/c = 0.60 and 0.80 induced by the 

LEX and wing vortices at the lower angles of 

attack and a transition to single-peaked 

distributions as the vortices directly interact and 

intertwine with each other at the higher angles of 

attack. In contrast, the porous LEX

configuration consistently promotes 

single-peaked pressure distributions with higher 

suction peak magnitudes at angles of attack 

above 10 degrees compared to the solid LEX 

configuration. These trends are consistent with 

the laser vapor screen flow visualization images 

at Mach = 0.85 that were previously shown in 

figure 104. As discussed reference 6, the 

porosity does not suppress the LEX vortex but, 

instead, promotes a weaker vortex and inhibits 

direct interaction with the wing vortical flow. 

The LEX vortex is sufficiently weak that its 

signature is not explicitly manifested in the 

pressure distributions. However, it induces a 

favorable effect on the wing flow field in a 

manner similar to a close-coupled canard 

(reference 3 and reference 14). Neither the 

solid LEX or porous LEX configurations show 

evidence of vortex breakdown effects up to the 

highest angle of attack of 28 degrees in 

figure 105. The wing-alone configuration

typically exhibits steeper gradients on either side 

of the suction peak and significantly higher 

suction peaks at x/c = 0.30. The latter effect is 

attributed to a longer “run length” from the wing 

apex for the generation of leading-edge vorticity. 

The primary benefit of the solid LEX and porous 

LEX at β = 0 degrees is to delay the onset of 

vortex breakdown effects, which begin to 

manifest in the wing-alone pressure distributions 

beginning at α = 28 degrees (see also figure 61).

The favorable effect of the LEX vortex on 

the stability of the wing vortical flow is more 

apparent in sideslip conditions. Figure 106

compares the pressure distributions obtained 

with the solid LEX, porous LEX, and wing 

alone at Mach = 0.50, α = 12, 16, 20, and 24 

degrees, and β = +/-4 degrees and β = +/-8 

degrees. Since the pressure taps were located 

exclusively on the right wing upper surface, 

negative sideslip angles (model apex yawed 

right) provide pressure distributions on the 

leeward wing, while positive sideslip angles 

(model apex yawed left) correspond to the 

windward wing. Positive sideslip angles 

decrease the direct interaction of the solid LEX 

and wing vortices, and the vortices are closer to 

the wing surface (reference 6). As a

consequence, the LEX and wing pressure 

signatures on the windward wing at β = +4 and 

+8 degrees are typically more distinct compared 

to the zero sideslip condition. Conversely, 

negative sideslip angles encourage a stronger 

direct interaction of the solid LEX and wing 

vortices, and the vortices are displaced away 

from the wing surface (reference 6). This 

decreases the LEX and wing vortex pressure

signatures on the leeward wing at β = -4 and 

-8 degrees The character of the solid LEX-wing 

pressure distributions is sensitive to the angle of 
30

attack and sideslip angle; however, there is no 

evidence of vortex breakdown even at the most 

extreme condition in figure 106 corresponding 

to α = 24 degrees and β = 8 degrees. The 

character of the porous LEX-wing pressure 

distributions is less sensitive to changes in the 

model attitude, although the windward wing 

exhibits markedly higher suction pressure peaks 

compared to the leeward wing. Similar to the 

solid LEX, vortex breakdown effects are absent 

from the porous LEX pressure distributions. In 

contrast, the wing-alone configuration shows 

initial signs of the advance of windward vortex 

breakdown over the wing beginning at 

α = 16 degrees and β = +8 degrees. Vortex 

breakdown effects dominate the wing-alone 

pressure distributions on the windward side at 

x/c = 0.60 and x/c = 0.80 at higher angles of 

attack where the pressure distributions become 

flatter and the overall suction pressure levels 

diminish.

Mach = 0.85

Figure 107 compares the upper surface static 

pressure coefficient distributions with the solid 

LEX and the LEX with full porosity at 

Mach = 0.85, angles of attack from 8 degrees to 

28 degrees in 2-degree increments, and 

β = 0 degrees. The wing-alone data are again 

shown for reference. Dual vortex suction 

pressure signatures can be discerned in the solid 

LEX pressure distributions at the lower angles of 

attack at x/c = 0.60 and x/c = 0.80. A single, 

broader vortex pressure signature with overall 

higher suction pressure level is apparent with the 

porous LEX. These trends are consistent with 

the vapor screen images in figure 104, which 

showed a broader wing vortex situated closer to 

the surface with the porous LEX. The porous 

LEX configuration exhibits consistently higher 

vortex-induced suction pressure levels in 

comparison to the solid LEX at the higher angles 

of attack. These results suggest that the flowthrough porosity provides an improved balance 

between the LEX and wing vortices at this Mach 

number. The wing vortex is thereby allowed to 

exert a greater influence on the high angle-ofattack flow field, unhindered by the direct 

interaction of a more dominant LEX vortical 

flow. The solid and porous LEX configurations 

retain distinct vortex pressure footprints at all 

measurement stations up to the highest angle of 

attack of 28 degrees, since the leading-edge 

vortices are stable at the higher angles of attack. 

In contrast, the shock-induced vortex breakdown 

that occurs with LEX off produces markedly 

flatter pressure distributions and lower suction 

pressure levels beginning at an angle of attack of 

24 degrees.

Figure 108 compares the pressure 

distributions obtained with the solid LEX, 

porous LEX, and wing-alone at Mach = 0.85, 

α = 12, 16, 20, and 24 degrees, and 

β = +/-4 degrees and β = +/-8 degrees. The 

trends at Mach = 0.85 are similar to those 

observed at Mach = 0.50 in figure 106, except 

the pressure distributions are typically broader 

and flatter at the higher Mach number. The 

windward wing pressure distributions on the 

solid LEX and porous LEX configurations at the 

higher angles of attack and x/c = 0.60 and 

x/c = 0.80 reveal subtle pressure peaks 

embedded in otherwise broad distributions. The

laser vapor screen flow visualization that was 

conducted in independent sideslip sweeps 

indicated these distributions were induced by 

stable vortex systems. It is noted that the porous 

LEX vortex pressure signature is still apparent at 

α = 24 degrees and β = +8 degrees; however, the 

suction pressure levels outboard of the suction 

peaks at x/c = 0.60 and x/c = 0.80 are 

measurably lower than the corresponding 

distributions obtained with the solid LEX. It is 

speculated that this effect is caused by a 

discontinuity in the leading-edge vortex feeding 

sheet such that the porous LEX vortex system 

“tears” away from the leading edge 

(reference 4). In contrast, the vapor screen flow 

visualization suggested the flatter pressure 

distributions on the wing-alone configuration 

beginning at α = 20 degrees and x/c = 0.80 and 

at α = 24 degrees and x/c = 0.60 and x/c = 0.80 

were associated with a rapid forward advance of 

a burst vortex system interacting with a normal 

shock wave.

Mach = 1.20

Figure 109 compares the upper surface static 

pressure coefficient distributions with the solid 
31

LEX and the porous LEX at Mach = 1.20, 

angles of attack from 8 degrees to 22 degrees in 

2-degree increments, and β = 0 degrees. The 

wing-alone data are included as a reference. The 

effect of LEX porosity on the wing surface 

pressure field is more subtle at this higher Mach 

number. In fact, the character of the pressure 

distributions and the overall suction pressure 

levels are similar for all configurations at 

Mach = 1.20. Vortex breakdown is not a factor 

on the wing-alone configuration up to the 

highest angle of attack of 22 degrees 

(reference 4). Consequently, the wing upper 

surface static pressure distributions are less 

sensitive to the presence of a vortex shed from 

the solid LEX or porous LEX. For similar 

reasons, the wing pressure distributions in 

sideslip shown in figure 110 at α = 12 and 

16 degrees and β = +/-4 degrees and 

β = +/-8 degrees are also less sensitive to the 

LEX vortex interaction with the wing flow field 

or to LEX porosity.

Force and Moment Measurements

Angle-of-Attack Sweeps at Mach = 0.50, 0.85, 

and 1.20 and β = 0 Degrees

Figure 111 through figure 113 compare the 

lift, drag, pitching moment, rolling moment, 

yawing moment, and side force coefficients 

obtained with the solid LEX and porous LEX in 

angle-of-attack sweeps at Mach = 0.50, 0.85, 

and 1.20, respectively, and β = 0 degrees 

(the wing-alone data are shown for reference).

Mach = 0.50

The porous LEX configuration exhibits 

similar lift coefficients at a given angle of attack

up to α ≅ 29 degrees, higher drag coefficients at 

a given lift coefficient, and reduced nose-up 

pitching moment coefficients compared to the 

solid LEX configuration at Mach = 0.50 

(figure 111). Although porosity weakens the 

LEX vortex, the overall lift is essentially 

unchanged because of the increased wing 

vortex-induced peak suction pressure levels that 

were apparent in the pressure distributions in 

figure 105. The higher drag coefficients are 

attributed to the flow through the porous surface, 

and the conjectured source of reduced pitch 

instability is the redistribution of lift from the 

LEX to the wing. The lift curve slopes are 

higher with the solid LEX and porous LEX 

configurations compared to the wing-alone, and 

the break in the wing-alone lift curve caused by 

the onset of vortex breakdown is absent in the 

presence of the LEX. It is noted, however, that 

the lift increase due to the LEX is typically less 

than the effect of a 15 percent area addition. 

This is not altogether surprising, since the 

65-degree leading-edge sweep of the isolated 

wing is conducive to the generation of a stable 

leading-edge vortex over a relatively wide range 

of angle of attack and correspondingly 

significant levels of vortex-induced lift 

(reference 4). The magnitude of the lateraldirectional aerodynamic coefficients are 

generally very small and do not exhibit 

significant aerodynamic asymmetries in the 

angle-of-attack sweeps at β = 0 degrees. 

However, the onset of vortex breakdown over 

the wing with LEX off at α = 28 degrees is 

slightly asymmetric, which causes the 

asymmetric break in the rolling moment 

coefficient curve in figure 111.

Mach = 0.85 

LEX porosity causes a slight increase in the 

lift coefficient at a given angle of attack 

compared to the solid LEX configuration at 

Mach = 0.85 as shown in figure 112. As 

previously noted, porosity weakens the LEX 

vortex. However, the pressure distributions in 

figure 107 indicated an overall increase in the 

wing vortex-induced suction pressure levels with 

the porous LEX. The abrupt drop-off in the lift 

coefficient with LEX off at α = 24 degrees is 

associated with an interaction of the wing vortex 

with a normal shock wave and an associated 

breakdown of the vortical flow. Consequently, 

it is concluded that the solid LEX and porous 

LEX are effective in eliminating the adverse 

effects of a normal shock wave over the wing. 

The weakening of the LEX vortex and the 

redistribution of the wing upper surface static 

pressure distributions are considered the primary 

sources of the nose-down increments in the 

pitching moment coefficients for the porous 

LEX configuration. Similar to the result 

obtained at Mach = 0.50 in figure 111, porosity 
32

typically increases the drag compared to the 

solid LEX due to the assumed jet-like flow 

through the porous surface. The shock-induced 

vortex breakdown over the wing-alone 

configuration is asymmetric, which promotes a 

significant asymmetry in the rolling moment 

coefficient curve at α = 24 degrees. 

Corresponding asymmetries occur in the yawing 

moment and side force coefficients. In contrast, 

the lateral-directional aerodynamic coefficients 

for the solid LEX and porous LEX 

configurations are relatively benign. 

Mach = 1.20

Increasing the Mach number to 1.20 limits 

the effect of porosity to a mitigation of the noseup pitching moment coefficients exhibited by 

the solid LEX configuration and a slight increase 

in the drag coefficient at a given lift coefficient 

(figure 113). The lift characteristics are similar 

for the solid LEX, porous LEX, and wing-alone

configuration through the range of angle of 

attack tested at this Mach number, since the 

LEX vortex-induced effects diminish at the 

higher Mach number (reference 4), and the 

effects of vortex breakdown are absent on all 

three configurations. This is consistent with the 

pressure distributions in figure 109, which 

showed similar vortex-induced suction pressure 

levels for the solid LEX, porous LEX, and wingalone configurations at this Mach number. No 

significant differences are observed in the 

lateral-directional aerodynamic coefficients.

Sideslip Sweeps at Mach = 0.50, 0.85, and 1.20

and Selected Angles of Attack

Figure 114 through figure 116 compare the 

six-component force and moment coefficients 

obtained with the solid LEX and porous LEX in 

sideslip sweeps at Mach = 0.50, 0.85, and 1.20, 

respectively, and at selected angles of attack (the 

wing-alone configuration is shown for 

reference).

 Mach = 0.50

 Figure 114 compares the longitudinal and 

lateral-directional aerodynamic coefficients 

obtained in sideslip sweeps at Mach = 0.50 and 

α = 12, 16, 20, and 24 degrees. The results at 

α = 12 degrees do not reveal any significant 

sensitivity of the lift, drag, and pitching moment 

coefficients to the sideslip angle for the solid 

LEX, porous LEX, or wing-alone 

configurations. Similarly, the variation of the 

lateral-directional aerodynamic coefficients with 

the sideslip angle is linear, and all configurations 

exhibit positive lateral-directional stability. The 

solid LEX configuration has a slightly higher 

level of lateral stability, whereas the wing-alone 

has marginally higher directional stability. 

Nonlinearities emerge in the longitudinal and 

lateral-directional characteristics at 

α = 16 degrees. For example, the wing-alone 

configuration exhibits breaks in the lift, drag, 

and pitching moment coefficients at moderate 

sideslip angles and corresponding unstable 

breaks in the rolling moment coefficient curve. 

This is caused by the onset of vortex breakdown 

effects on the windward wing, which were 

inferred from the pressure distributions in 

figure 89 and figure 106. The porous LEX 

configuration also shows some sensitivity to the 

sideslip angle, which is manifested by a slight 

decrease in the lift and drag coefficients at the 

extremes of the sideslip angle range and a 

corresponding decrease in the lateral stability 

compared to the solid LEX configuration. There 

is no indication of the onset of vortex 

breakdown effects. Instead, it is speculated that 

the pressure differential between the windward 

and leeward wings is not as great with the 

weaker porous LEX vortices compared to the 

solid LEX. This trend was implied in the 

pressure distributions in figure 71 and figure 78, 

which showed larger differences in the 

windward and leeward wing suction pressure 

levels in the presence of the solid LEX. These 

effects are magnified at α = 20 and 24 degrees. 

Mach = 0.85

 Figure 115 compares the longitudinal and 

lateral-directional aerodynamic coefficients 

obtained in sideslip sweeps at Mach = 0.85 and

α = 12, 16, 20, and 24 degrees. Similar to the 

results obtained at Mach = 0.50 in figure 114, 

the data at Mach = 0.85 and α = 12 degrees do 

not reveal any significant sensitivity of the lift, 

drag, and pitching moment coefficients to the 

sideslip angle for the solid LEX, porous LEX, or 

wing-alone configurations. Furthermore, the 

variation of the lateral-directional aerodynamic 

coefficients with the sideslip angle is linear, and 
33

all configurations exhibit positive lateraldirectional stability. Shock-induced vortex 

breakdown effects on the windward wing with 

LEX off are inferred at α = 16 degrees from the 

abrupt decrease in the lift and drag coefficients, 

nose-up pitching moment coefficient increments, 

and the unstable breaks in the rolling moment 

coefficient curve at the extreme ends of the 

sideslip sweep. The lift, drag, and pitching 

moment coefficients for the solid LEX 

configuration are generally insensitive to 

variation in the sideslip angle, whereas the 

porous LEX configuration displays a decrease in 

the lift and drag coefficients as the model 

traverses through the higher (in absolute value) 

sideslip angles. However, both configurations 

have comparable levels of static 

lateral-directional stability. Similar trends are 

observed for the solid LEX and porous LEX at 

α = 20 degrees, whereas the wing-alone 

configuration exhibits more significant 

aerodynamic nonlinearities as the effects of the 

normal shock wave are magnified. The primary 

difference in trends at α = 24 degrees is the 

apparent onset of shock-induced vortex 

breakdown on the porous LEX configuration at 

the higher values of sideslip angles, which is 

manifested by the aerodynamic nonlinearities at 

the extremes in the sideslip sweep. This effect 

was also inferred from the porous LEX pressure 

distributions in figure 79. In addition, LVS flow 

visualization (not shown) indicated an instability 

of the windward wing vortex system at the 

sideslip extremes with the porous LEX.

Mach = 1.20

 Figure 116 compares the longitudinal and 

lateral-directional aerodynamic coefficients 

obtained in sideslip sweeps at Mach = 1.20 and 

α = 12 and 16 degrees. There are no significant 

differences in the six-component aerodynamic 

force and moment coefficient trends with the 

sideslip angle at these angles of attack, since 

shock-induced vortex breakdown effects are 

absent at this Mach number. The wing-alone

configuration has slightly higher levels of 

directional stability at both angles of attack, 

which may be caused by different 

vortex-induced sidewash effects at the vertical 

tail.

Full Versus Partial LEX Porosity

Pressure Distributions

Porosity Levels 1, 2 and 3

Figure 117 compares the upper surface static 

pressure distributions obtained at Mach = 0.50 

and α = 8, 12, 16, 20, 24, and 28 degrees with 

porosity levels 1, 2, and 3. Porosity level 1 

corresponds to the uniform 14.8 percent

porosity; porosity levels 2 and 3 were obtained 

by compartmentalizing the porosity to the 

regions aft and forward of the LEX planform 

break, respectively (9.4 percent and 5.4 percent 

porosity). It is not known if the 

compartmentalized porosity promoted multiple 

vortex development from the LEX, since LVS

flow visualization was not performed on the 

partial porosity configurations. The differences 

in the respective pressure distributions are 

nominal, however, and the character of the 

pressure distributions appears independent of the 

porosity level. Porosity level 1 promotes 

slightly higher vortex suction pressure peaks that 

are typically farther outboard, whereas porosity 

levels 2 and 3 generally result in higher overall 

suction pressure levels along the inboard region 

of the wing.

The trends are similar, albeit more subtle, at 

Mach = 0.85 (figure 118) up to approximately 

24 degrees angle of attack. At α = 28 degrees, 

however, there is a marked decline and 

flattening of the pressure distributions at 

x/c = 0.60 and x/c = 0.80 with porosity level 3, 

which suggests this configuration is prone to 

shock-induced vortex breakdown effects.

Porosity Levels 1, 4 and 5

Figure 119 compares the upper surface static 

pressure distributions obtained at Mach = 0.50 

and α = 8, 12, 16, 18, 24, and 28 degrees with

porosity levels 1, 4, and 5. Comparisons at 

Mach = 0.85 and α = 8, 12, 18, 20, 24, and 28 

degrees are shown in figure 120. Valid data 

points were not acquired at α = 20 degrees 

and α = 16 degrees at Mach = 0.50 and 0.85, 

respectively, for the configuration with porosity 

level 5. Porosity level 1 corresponds to the 

uniform 14.8 percent porosity; porosity levels 4
34

and 5 were obtained by applying porosity to a 

band along the leading-edge region or to the 

region inboard of this band, respectively 

(6.1 percent and 8.7 percent porosity). LVS

flow visualization was not performed on the 

partial porosity configurations.

The configuration with full porosity typically 

promotes higher suction pressure peaks at 

Mach = 0.50 in figure 119, although the 

differences in the pressure distributions at the 

three measurement stations diminish 

considerably at the higher angles of attack. The 

latter trend is also apparent in the pressure 

distributions at Mach = 0.85 in figure 120. The 

pressure distributions obtained at the subsonic 

and transonic Mach numbers suggest that a 

porous band about the leading edge may be as 

effective as full porosity applied to the entire 

LEX surface.

Force and Moment Measurements

Porosity Levels 1, 2 and 3

Figure 121 compares the longitudinal 

aerodynamic coefficients obtained with porosity 

levels 1, 2, and 3 in angle-of-attack sweeps at 

Mach = 0.50. There are no significant 

differences in the lift and drag coefficients 

obtained with the three porosity levels. The 

configuration with porosity level 3 exhibits more 

nose-up pitching moment coefficients at 

moderate and high lift coefficients. There are no 

aerodynamic discontinuities associated with the 

onset of vortex breakdown over the wing up to 

the highest angle of attack of approximately 29 

degrees at Mach = 0.50. These results are 

consistent with the pressure distribution 

comparisons in figure 117.

The longitudinal aerodynamic characteristics 

are similar for all three porosity levels at 

Mach = 0.85 in figure 122 up to an angle of 

attack of approximately 26 degrees. At higher 

angles of attack, the configuration with porosity 

level 3 shows an abrupt drop-off in the lift 

coefficient, increased drag at a given lift, and 

nose-up pitching moment increments that are

assumed to be caused by shock-induced vortex 

breakdown. This effect was also inferred from 

the pressure distributions in figure 118. 

Porosity Levels 1, 4 and 5

Figure 123 compares the longitudinal 

aerodynamic coefficients obtained with porosity 

levels 1, 4, and 5 in angle-of-attack sweeps at 

Mach = 0.50. No notable differences are 

apparent in the lift, drag, and pitching moment

coefficients obtained with the three porosity 

levels at this Mach number. At Mach = 0.85

(figure 124), however, the configuration with 

porosity level 4 displays marked aerodynamic 

discontinuities in all three longitudinal force and 

moment coefficients as the angle of attack 

increases beyond α = 28 degrees. The character 

of these nonlinearities is consistent with 

shock-induced vortex breakdown at this 

transonic Mach number. The pressure 

distributions in figure 120 did not reveal this 

effect, since data comparisons for all three 

porosity levels at the same angle of attack were 

not available. The pressure distributions and 

force and moment characteristics suggest that 

the porosity level may be tailored to obtain a 

desired aerodynamic effect (for example, 

reduced nose-up pitching moments). However, 

compartmentalizing and reducing the level of 

LEX porosity renders the configuration 

susceptible to aerodynamic discontinuities 

associated with vortex breakdown.

Comparison of Vertical Tail

Arrangements

Pressure Distributions

Mach = 0.50, β = 0 Degrees

Figure 125 compares the wing upper surface 

static pressure distributions obtained with the 

solid LEX and the centerline tail, twin canted 

tails, twin uncanted tails, and tail off at 

Mach = 0.50 and α = 8, 12, 16, 20, 24, and 28 

degrees. Data points corresponding to the 

tail-off configuration were typically acquired at 

off-setpoint conditions for angles of attack 

greater than 24 degrees. As a result, the tail-off 

configuration is not included in the data 

comparisons at α = 28 degrees. As will be 

discussed later, however, the centerline tail and 

tail-off configurations yielded similar pressure 
35

distributions at all angles of attack and zero 

sideslip conditions. It is noted that the junction 

of the canted and uncanted tail leading edges 

with the wing upper surface is just aft of the 

80 percent wing chord station. At α = 8 degrees, 

the subtle pressure peak induced by the LEX 

vortex at x/c = 0.80 for the centerline tail and 

tail-off configurations is absent with the twin 

canted and uncanted tails. The pressure 

distributions are otherwise similar in all respects 

at the three measurement stations. Limited LVS 

flow visualization obtained in reference 5

indicates that the canted and uncanted tails are in 

the path of the LEX vortex but do not promote 

early vortex breakdown at this relatively low 

angle of attack. Instead, the vortex is able to 

navigate about the tail without bursting. Similar 

effects have been observed at supersonic speeds 

in reference 6. A similar trend is observed at 

α = 12 degrees. In addition, the wing pressure 

signature at x/c = 0.80 is reduced in the presence 

of the twin uncanted tail. At α = 16 degrees, the 

canted and uncanted tails promote a significant 

decrease in the overall suction pressure level at 

x/c = 0.80 compared to the centerline tail and 

tail-off configurations, although it is not possible 

to infer from the pressure distributions if the tail 

interaction with the wing flow field causes 

vortex breakdown. The most significant effect 

is caused by the canted tail at this angle of 

attack. There is also a slight decrease in the 

suction pressure levels at x/c = 0.60 with both 

wing-mounted tail arrangements, which is 

indicative of a forward propagation of the tail 

effect on the vortex flows. Although the overall 

suction pressure levels differ, the character of

the pressure distributions at x/c = 0.80 is similar 

for all tail configurations, which suggests that 

the vortices are displaced away from the wing 

surface in the presence of the wing-mounted 

tails. Evidence that the canted tail causes a more 

significant intrusive effect on the vortical flows 

is provided in the pressure data at 

α = 20 degrees and at x/c = 0.60 and x/c = 0.80, 

where the overall suction pressure levels are 

significantly reduced compared to the centerline 

tail and tail-off configurations. The uncanted 

tail has a similar, but less dramatic, effect on the 

pressure distributions. These effects are 

magnified at α = 24 and 28 degrees, and the 

pressure distributions at x/c = 0.30 also display a 

sensitivity to the presence of the wing-mounted 

tails. The greater influence of the canted tail on 

the wing pressure distributions suggests that this 

tail orientation places it more directly in the path 

of the interacting solid LEX and wing vortices. 

Despite the significant decline in the overall 

suction pressure levels with the canted and 

uncanted tails, the pressure distributions 

maintain discernible pressure coefficient 

maxima and minima, which are consistent with 

unburst vortices and regions of flow 

reattachment, respectively. Consequently, it is 

speculated that the primary effect of the canted 

and uncanted tails is an upward displacement of 

the LEX and wing vortices and a subsequent 

decrease in their respective surface pressure 

signatures. In contrast, there is no significant

difference in the pressure distributions obtained 

on the centerline tail and tail-off configurations

at any angle of attack, which suggests the 

centerline tail has a minimal effect on the vortex 

flow field at this zero-sideslip condition.

Mach = 0.85, β = 0 Degrees

Figure 126 compares the wing upper surface 

static pressure distributions obtained with the 

solid LEX and the centerline tail, twin canted 

tails, twin uncanted tails, and tail off at 

Mach = 0.85 and α = 8, 12, 16, 20, 24, and 28 

degrees. The discussion of the tail effects at 

Mach = 0.50 is mostly applicable to the data 

obtained at Mach = 0.85. The most notable 

exception at this transonic Mach number is the 

tail effects on the pressure distributions are 

largely confined to the aft measurement station 

up to angles of attack of 24 degrees. At 

α = 28 degrees, only the canted tail has an 

upstream effect that propagates forward to 

x/c = 0.60. There is insufficient data to 

determine if vortex breakdown occurs in the 

presence of the wing-mounted tails. The LVS

flow visualization results obtained on the solid 

LEX-wing configuration with centerline tail in 

figure 104 revealed a stable vortex system up to 

at least an angle of attack of 28 degrees at 

Mach = 0.85. 
36

Mach = 0.50, β = +/-4 and +/-8 Degrees

Figure 127 compares the wing upper surface

static pressure distributions obtained with the 

centerline tail, twin canted tails, twin uncanted 

tails, and tail off at Mach = 0.50, α = 12, 16, 20, 

and 24 degrees, and β = +/-4 degrees and 

+/-8 degrees. Intrusive effects of the twin tails 

are largely confined to the pressure distributions 

at x/c = 0.80, except at the highest angle of 

attack of 24 degrees. In general, the windward 

and leeward wing pressure distributions are 

clustered in two distinct groups at this 

measurement station corresponding to (1) the 

centerline tail and tail off and (2) the canted and 

uncanted tails. The differences in the pressure 

distributions at α = 12 degrees are relatively 

small. The wing-mounted tails generally cause a 

slight decrease in the overall suction pressure 

levels on the windward and leeward wings at 

x/c = 0.80. This effect is magnified at α = 16, 

20, and 24 degrees and is attributed to an 

upward displacement of the vortices from the 

wing surface in the presence of the twin tails. 

The uncanted tails also promote an overall 

decrease in the suction pressure levels on the 

windward wing at α = 24 degrees, x/c = 0.60

and 0.80, and β = +4 and +8 degrees, although 

there is insufficient data to infer the onset of 

vortex breakdown caused by an interaction of 

the vortices with the wing-mounted tails.

Mach = 0.85, β = +/-4 and +/-8 Degrees

The tail effects on the windward and leeward 

wing pressure distributions at Mach = 0.85 in 

figure 128 are largely confined to the aft 

measurement station at angles of attack up to 

20 degrees, where both wing-mounted tail 

configurations cause a reduction in the overall 

suction pressure levels compared to the 

centerline tail and tail-off configurations. 

Interestingly, the latter two configurations

exhibit relatively flat pressure distributions on 

the windward wing at α = 20 degrees, 

β = +8 degrees, and x/c = 0.80, although there is 

no evidence to confirm that vortex breakdown 

has occurred over the wing. In fact, the 

character of the windward wing pressure 

distributions with the canted and uncanted tails

is consistent with stable vortices, since it is 

similar to the pressure signatures corresponding 

to the centerline tail and tail-off configurations.

The presence of the centerline tail is also 

inferred from the leeward wing pressure 

distributions at x/c = 0.80 by a slight decrease in 

the overall suction pressure level compared to 

the tail-off case. Similar trends are observed at 

α = 24 degrees, except the upstream influence of 

the uncanted tail is observed in the windward 

wing pressure distributions at x/c = 0.60.

Force and Moment Measurements

Angle-of-Attack Sweeps at Mach = 0.50 and 

0.85 and β = 0 Degrees

Figure 129 and figure 130 compare the lift,

drag, pitching moment, rolling moment, yawing 

moment, and side force coefficients obtained in 

angle of attack sweeps at zero sideslip with the 

solid LEX and the centerline tail, twin canted 

tails, twin uncanted tails, and tail off at 

Mach = 0.50 and 0.85, respectively.

Mach = 0.50

Both wing-mounted tail configurations result 

in significant loss of lift, increased drag at a 

given lift, and large nose-up pitching moment 

increments compared to the centerline tail and 

tail-off configurations at Mach = 0.50 

(figure 129) beginning at approximately 

12 degrees angle of attack. In addition, the 

longitudinal aerodynamic characteristics are 

nonlinear, which reflects the varying manner in 

which the LEX and wing vortices interact with 

the canted and uncanted tails through the range 

of angle of attack. The pressure distributions in 

figure 125 were inconclusive regarding the 

occurrence of vortex breakdown in the presence 

of the wing-mounted tails. It is assumed the 

longitudinal aerodynamic data trends are 

associated with changes in which the LEX and 

wing vortices interact with each other and with 

the canted or uncanted tails and, also, with an 

upward displacement of the vortices from the 

wing surface. Aerodynamic asymmetries are 

evident in the lateral-directional aerodynamic 

characteristics of the uncanted tail configuration 

at α = 26 degrees. Asymmetries in the yawing 

moment and side force coefficients are also 
37

present at α = 28 degrees for the canted tail 

configuration.

Mach = 0.85

Similar nonlinear characteristics caused by 

the wing-mounted tails are present in the 

longitudinal aerodynamic characteristics at 

Mach = 0.85 in figure 130. There are no 

significant lateral-directional aerodynamic 

asymmetries for any configuration up to 

α = 28 degrees. However, data acquired at 

α = 30 degrees for the tail-off configuration 

indicates a large roll-off which is concurrent 

with an abrupt loss of lift. 

Sideslip Sweeps at Mach = 0.50 and 0.85 and 

Selected Angles of Attack

Figure 131 and figure 132 compare the sixcomponent force and moment coefficients 

obtained at Mach = 0.50 and 0.85, respectively, 

in sideslip sweeps at selected angles of attack 

with the centerline tail, twin canted tails, twin 

uncanted tails, and tail off.

Mach = 0.50

The canted and uncanted tail configurations 

exhibit nonlinear variations in the lift, drag, and 

pitching moment coefficients with the sideslip 

angle at Mach = 0.50 and α = 12 degrees in 

figure 131. The uncanted tail configuration 

regains some lift on either side of β = 0 degrees, 

which is presumably the result of the LEX 

vortex navigating away from the vertical tail. 

Conversely, the canted tail configuration 

displays a lift loss on either side of 

β = 0 degrees, since the LEX vortex is more 

aligned with the canted tail position in sideslip. 

Despite the nonlinearities in the longitudinal 

aerodynamic characteristics, the rolling moment 

curves for both wing-mounted tail 

configurations are linear and stable across the 

range of sideslip angle, and all four 

configurations exhibit similar levels of static 

lateral stability. Nonlinearities are present in the 

yawing moment coefficient curves, which reflect 

the varying interaction of the vortices with the 

wing-mounted tails. Although the tail volume is 

essentially the same for all tail configurations, 

the centerline tail configuration has the highest 

level of static directional stability, since the 

single tail is not prone to direct interaction with 

the vortical flows. The configuration without 

tail is statically unstable in yaw. The nonlinear 

longitudinal aerodynamic coefficient data trends 

with the wing-mounted tails are somewhat 

mitigated at α = 16 degrees, and all 

configurations continue to display relatively 

linear, stable variations of the rolling moment 

coefficient with sideslip. The lateral stability of 

the centerline tail configuration is essentially 

unchanged compared to α = 12 degrees, whereas 

the canted and uncanted tail configurations 

exhibit reduced or neutral directional stability at 

this higher angle of attack. Longitudinal 

aerodynamic nonlinearities on the wingmounted tail configurations re-emerge at 

α = 20 and 24 degrees, particularly with the 

uncanted tails. The interaction of the LEX and 

wing vortices with the wing-mounted tails also 

promotes nonlinearities in the rolling moment 

and yawing moment coefficient curves in 

contrast to the linear behavior exhibited by the 

centerline tail and tail-off configurations.

Mach = 0.85

The nonlinear longitudinal aerodynamic 

coefficient data trends associated with the twin 

canted and uncanted tails observed at 

Mach = 0.50 are also apparent at Mach = 0.85 in

figure 132. The discontinuities in the 

longitudinal aerodynamic coefficients at 

α = 12 degrees near β = 0 degrees with the 

uncanted tails is the result of inadvertent data 

acquisition at angles of attack below the desired 

setpoint. In a typical β-sweep, the model 

sideslip angle was varied from 0 degrees to the 

maximum positive sideslip angle, then returned 

to a nominal β = 0 degrees condition. The 

β-sweep was then completed by varying the 

sideslip angle from 0 degrees to the maximum 

negative sideslip angle. At each sideslip angle, 

the model attitude was manually adjusted in an 

attempt to maintain the angle of attack within 

approximately 0.05 degrees of the desired 

setpoint. This procedure was occasionally 

violated, however, during the transition from the 

positive sideslip to negative sideslip portions of 

the β-sweep.
38

The longitudinal aerodynamic nonlinearities

with the uncanted tails are less pronounced at

Mach = 0.85 and the higher angles of attack 

(α = 20 and 24 degrees), which suggests the 

vortex-tail interaction is sensitive to the Mach 

number. The variation of the rolling moment 

coefficient with sideslip is linear and stable at 

Mach = 0.85 up to α = 24 degrees for both twintail configurations, in contrast to the nonlinear 

behavior exhibited at Mach = 0.50 in figure 31. 

The yawing moment curves in figure 132 at 

α = 20 and 24 degrees are nonlinear with the 

wing-mounted tails. However, the directional 

stability levels are comparable to, or higher than, 

the corresponding results obtained with the 

centerline tail configuration at this Mach 

number.

Comparison of Solid LEX and Canard

Pressure Distributions

Mach = 0.50

Figure 133 compares the upper surface static 

pressure coefficient distributions with the solid 

LEX and the canard at Mach = 0.50 and angles 

of attack from 8 degrees to 28 degrees in 

2-degree increments. The wing-alone pressure 

distributions are included in each figure for 

reference. All configurations featured a 

centerline vertical tail, and all plots correspond 

to a nominal zero-sideslip condition. The canard 

downwash field decreases the wing vortexinduced suction pressures and displaces the 

suction peak outboard at x/c = 0.30, which is 

within the span of the canard. This effect is 

manifested at all angles of attack from 8 to 28 

degrees and is indicative of a smaller, weaker

vortex that is closer to the wing leading edge. 

The solid LEX promotes a similar effect at this 

pressure measurement station. The LEX vortex 

pressure signature is apparent at x/c = 0.30, 

whereas the canard vortex footprint is absent 

since it is situated higher above the wing surface 

(reference 14). A primary difference between 

the solid LEX-wing and canard-wing flow fields 

is the level of interaction with the wing flow 

field. The LEX vortex is highly-coupled to the 

wing flow field such that the LEX and wing 

vortices directly interact with each other. 

Although LVS flow visualization was not 

conducted on the canard-wing configuration in 

the current investigation, flow visualization 

results obtained on the same model in 

reference 14 indicate the relative spacing of the 

canard and wing vortices does not encourage 

intertwining of the vortices. Both the LEX and 

canard promote a favorable flow-field 

interference with the wing that results in an 

effective increase in the wing leading-edge 

sweep (reference 16), stabilization of the wing 

vortex, and mitigation of transonic shockinduced flow separation and vortex breakdown. 

The canard-induced upwash field increases the 

effective angle of attack at the wing leading 

edge at x/c = 0.60 and x/c = 0.80, which are 

outside the canard span. As a result, the wing 

vortex suction peaks at these measurement 

stations increase to levels comparable to those 

obtained on the isolated wing. At α = 8 to 16 

degrees, the wing pressure distributions in the 

presence of the canard at x/c = 0.60 and 

x/c = 0.80 exhibit a single suction pressure peak 

induced by the wing vortex that is situated close 

to the leading edge. The wing vortex pressure 

signature is of comparable magnitude in the 

presence of the LEX, but is situated farther 

inboard. Furthermore, the pressure signature of 

the LEX vortex is also apparent at these angles 

of attack. In general, the overall suction 

pressure levels at these measurement stations are 

higher on the LEX-wing configuration. As the 

LEX and wing vortices directly interact at the 

higher angles of attack, the resultant vortex 

system typically induces higher suction pressure 

peaks that are situated farther inboard and 

overall suction pressures that exceed those 

obtained with the canard. The pressure 

distributions obtained on the LEX-wing and 

canard-wing configurations are consistent with 

stable vortices up to the highest angle of attack 

of 28 degrees shown in figure 133.

Figure 134 compares the wing upper surface 

static pressure distributions obtained with the 

solid LEX, canard, and wing-alone at 

Mach = 0.50, α = 12, 16, 20, and 24 degrees, 

and β = +/-4 degrees and +/-8 degrees. The 

primary difference in the windward wing 
39

(β = +4 and +8 degrees) pressure distributions in 

the presence of the LEX and canard at 

α = 12 and 16 degrees is the pressure signature 

induced by the LEX vortex and the associated 

higher suction pressure levels along the inboard 

portion of the wing. On the leeward wing 

(β = -4 and -8 degrees), the wing vortex is 

situated closer to the leading edge, and the 

overall suction pressure levels are typically less 

in the presence of the canard at x/c = 0.60 and

x/c = 0.80. These effects are magnified at 

α = 20 and 24 degrees. In general, the pressure 

data suggest the positional changes of the wing 

vortex due to changes in the sideslip angle are 

reduced in the presence of the canard. The 

overall suction pressure levels on the windward 

and leeward wings are generally higher on the 

LEX-wing configuration because of the 

stronger, more direct interaction of the LEX and 

wing vortices. Both the LEX and canard 

eliminate the effects of windward wing vortex 

breakdown that are apparent in the wing-alone 

pressure distributions at α = 24 degrees.

Mach = 0.85

Figure 135 compares the upper surface static 

pressure coefficient distributions with the solid 

LEX, canard, and wing-alone at Mach = 0.85, 

angles of attack from 8 degrees to 28 degrees in 

2-degree increments, and β = 0 degrees. The 

pressure distributions obtained on these three 

configurations at Mach = 0.85, α = 12, 16, 20, 

and 24 degrees, and β = +/-4 degrees and 

+/-8 degrees are compared in figure 136. The 

trends at this transonic Mach number are similar 

to those observed at Mach = 0.50 in figure 133

and figure 134. In general, the differences in 

the overall suction pressure levels between the 

LEX and canard configurations diminish

significantly at the higher Mach number, and the 

pressure distributions are typically broader with 

less pronounced suction peaks. The shockinduced vortex breakdown effects that occur on 

the wing-alone at α = 24, 26, and 28 degrees are 

mitigated by the presence of the LEX and 

canard. 

Force and Moment Measurements

Angle-of-Attack Sweeps at Mach = 0.50 and 

0.85 and β = 0 Degrees

Mach = 0.50

Figure 137 compares the lift, drag, pitching 

moment, rolling moment, yawing moment, and 

side force coefficients obtained with the solid 

LEX and canard in angle-of-attack sweeps at 

Mach = 0.50 and β = 0 degrees (the wing-alone 

configuration is shown for reference). The solid 

LEX configuration exhibits slightly higher lift 

coefficients at angles of attack greater than 

approximately 12 degrees, less drag at a given 

lift, and less nose-up pitching moment compared 

to the canard. These results are consistent with 

the pressure distributions in figure 133, which 

generally showed higher overall suction pressure 

levels on the wing in the presence of the solid 

LEX. The solid LEX and canard configurations 

eliminate the lift curve break associated with 

incipient vortex breakdown exhibited by the 

wing-alone configuration. However, neither the 

LEX-wing or canard-wing flow-field interaction 

results in a favorable lift synergism, since the 

percentage increase in lift is consistently less 

than the area increase due to the LEX or canard. 

The canard and wing-alone configurations

develop lateral-directional aerodynamic 

asymmetries at the higher angles of attack, 

although the magnitude of these asymmetries is 

relatively small.

Mach = 0.85

In contrast to the results obtained at 

Mach = 0.50, the canard configuration develops 

slightly higher lift at angles of attack greater 

than approximately 12 degrees and less drag at a 

given lift at Mach = 0.85 as shown in figure 138. 

The canard configuration develops more nose-up 

pitching moment coefficients compared to the 

solid LEX configuration, which is similar to the 

results observed at Mach = 0.50. In addition, the 

variation of the pitching moment coefficient 

with the lift coefficient is more nonlinear. The

abrupt breaks in the lift, drag, and pitching 

moment curves and the large lateral-directional 

aerodynamic asymmetries caused by shockinduced vortex breakdown on the wing-alone 
40

configuration are absent on the solid LEX and 

canard configurations.

Sideslip Sweeps at Mach = 0.50 and 0.85 and 

Selected Angles of Attack

 Mach = 0.50

 Figure 139 compares the longitudinal and 

lateral-directional aerodynamic coefficients 

obtained in sideslip sweeps at Mach = 0.50

and α = 12, 16, 20, and 24 degrees on the solid 

LEX, canard, and wing-alone configurations. 

The discontinuity in the longitudinal 

aerodynamic coefficients with canard at 

α = 20 degrees near β = 0 degrees is caused by 

data acquisition at angles of attack below the 

desired setpoint as previously described in 

figure 132. The lift, drag, and pitching moment 

coefficients obtained on the solid LEX and 

canard configurations are generally insensitive 

to changes in the sideslip and do not exhibit the 

nonlinear breaks displayed by the wing-alone 

configuration at α = 16, 20, and 24 degrees. 

Similarly, the LEX and canard configurations 

have linear and stable variations of the rolling 

moment coefficient with the sideslip angle, in 

contrast to the unstable breaks in the rolling 

moment coefficient curves for the wing alone. 

Slightly higher lateral stability is obtained with 

the LEX, which is consistent with the higher 

overall suction pressures on the windward wing 

in figure 134. All three configurations are 

directionally stable, although the canard 

configuration has the lowest level of directional 

stability at each angle of attack.

Mach = 0.85

 The trends in the longitudinal and lateraldirectional aerodynamic coefficients obtained in 

sideslip sweeps at Mach = 0.85 in figure 140 are 

similar to those described in figure 139 at 

Mach = 0.50. Discontinuities in the data about 

β = 0 degrees are again attributed to data 

acquisition at angles of attack slightly different 

from the desired setpoints. The LEX and canard 

configurations are not subject to the 

nonlinearities in the lift, drag, pitching moment, 

and rolling moment coefficients curves due to 

shock-induced vortex breakdown that occur on 

the wing alone. All three configurations exhibit 

nonlinearities in the yawing moment and side 

force coefficient curves at α = 24 degrees, which 

reflects the different interactions of the 

separated/vortex flow fields with the centerline 

tail. The solid LEX-wing configuration is 

essentially neutrally-stable in yaw at this angle 

of attack, whereas the canard-wing configuration 

is slightly unstable.

Concluding Remarks

The effects of passive surface porosity on 

vortex flow interactions about a general research 

fighter configuration were investigated at 

subsonic and transonic speeds. Flow-through 

porosity was applied to a wing leading-edge 

extension mounted to a 65-degree cropped delta 

wing model as a longitudinal instability 

mitigation technique at high angles of attack. 

The porosity of the LEX was also 

compartmentalized to determine the sensitivity 

of the vortex-dominated aerodynamics to the 

location and level of porosity. Test data were 

obtained with LEX on and off in the presence of 

a centerline vertical tail and twin, wing-mounted 

vertical fins having 0-degree and 30-degree cant 

angles to quantify the sensitivity of the vortex 

flow aerodynamics to tail placement and 

orientation. A close-coupled canard was also 

tested as an alternative to the wing LEX as a 

high angle-of-attack passive flow control device. 

Wing upper surface static pressure distributions 

and six-component forces and moments were 

obtained at Mach numbers of 0.50, 0.85, and 

1.20, unit Reynolds number of 2.5 million, 

angles of attack up to approximately 30 degrees,

and angles of sideslip to +/-8 degrees. The 

off-surface flow field was visualized in cross 

planes on selected configurations using a laser 

vapor screen flow visualization technique. A 

subset of data was obtained on selected 

configurations at Mach = 0.50 to identify 

Reynolds number effects on the surface pressure 

distributions and force and moment 

characteristics. Within-test data repeatability 

was assessed. Comparisons were also made of 

the pressure distributions and aerodynamic force 

and moment coefficients obtained on several 

configurations tested at Mach = 0.50 in the

NASA LaRC 8-Foot TPT and in a previous test
41

using the same model in the NASA LaRC 

7- by 10-Foot High Speed Tunnel. The test 

results were insensitive to the Reynolds number 

within the scope of the present experiment. 

Within-test and tunnel-to-tunnel data 

repeatability were satisfactory. Porosity 

decreased the vorticity shed from the LEX, 

which weakened the LEX vortex and altered the 

global interactions of the LEX and wing vortices 

at high angles of attack. The desired nose-down 

pitching moment increments were concurrent 

with similar maximum lift, higher drag-due-tolift, and reduced lateral-directional stability. 

Compartmentalizing the LEX porosity did not 

offer any advantages over the LEX with full 

porosity. The vortex-dominated aerodynamics 

were sensitive to the vertical tail arrangement, 

and the centerline tail configuration yielded the 

most desirable longitudinal and lateraldirectional aerodynamic characteristics at all 

Mach numbers tested. The wing upper surface 

static pressure distributions obtained with the 

canard and the solid LEX were distinct. At zero 

sideslip, the solid LEX typically exhibited 

slightly higher maximum lift and less drag and 

nose-up pitching moment at a given lift. 

References 

1. Bauer, S. X. S. and Hemsch, M. J., 

“Alleviation of Side Force on TangentOgive Forebodies Using Passive Porosity,” 

AIAA-92-2711, June 1992. 

2. Bauer, S. X. S. and Hernandez, G., 

“Reduction to Cross-Flow Shock-Induced 

Separation with a Porous Cavity at 

Supersonic Speeds,” AIAA-88-2567, June 

1988. 

3. 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, Oct. 1986, pp. 17-19. 

4. 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. 

5. Erickson, G. E. and Inenaga, A. S., “FiberOptic-Based Laser Vapor Screen Flow 

Visualization for Aerodynamic Research in 

Larger Scale Subsonic and Transonic Wind 

Tunnels,” NASA TM 4514, 1994. 

6. 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, October 

2007. 

7. Braslow, A. L., Hicks, R. M., and Harris, R. 

V. Jr., “Use of Grit-Type Boundary-Layer 

Transition Strips on Wind-Tunnel Models,” 

NASA TN D-3579, 1966. 

8. Hall, R. M., Erickson, G. E., and Fox, C. H. 

Jr., “Evaluation of Gritting Strategies for 

High Angle of Attack Using Wind Tunnel 

and Flight Test Data for the F/A-18,” NASA 

TP-1998-207670, May 1998. 

9. Brooks, C. W., Harris, C. D., and Reagon, P. 

G., “The NASA Langley 8-Foot Transonic 

Tunnel Calibration,” NASA TP 3437, 1994. 

10. Jackson, C. M. Jr., Corlett, W. A., and 

Monta,W. J., “Description and Calibration 

of the Langley Unitary Plan Wind Tunnel,” 

NASA TP 1905, November 1981. 

11. Gillis, C. L., Polhamus, E. C., and Gray, J. 

L. Jr., “Charts for Determining Jet-Boundary 

Corrections for Complete Models in the 

7- by 10-Foot Closed Rectangular Wind 

Tunnels,” NACA WR L-123, 1945. 

12. Herriot, J. G., ”Blockage Corrections for 

Three-Dimensional Flow Closed-Throat 

Wind Tunnels, With Consideration of the 

Effect of Compressibility,” NASA Rep. 

995, 1950. (Supersedes NACA RM A7B28.) 

13. Graham, A. B., “Data Engineering Scripting 

Language (DESL), ViGYAN, Inc., 

Hampton, VA. 

14. Erickson, G. E., Schreiner, J. A., and 

Rogers, L. W., “Canard-Wing Vortex 

Interactions at Subsonic Through Supersonic 

Speeds,” AIAA-90-2814, August 1990. 

15. Erickson, G. E., “Wind Tunnel Investigation 

of Vortex Flows on F/A-18 at Subsonic 

Through Transonic Speeds,” NASA 

TP-3111, 1991. 

16. Hemsch, M. J. and Luckring, J. M., 

“Connection Between Leading-Edge Vortex 

Lift and Vortex Strength for Delta Wings,” 

J. Aircr., vol. 27, no. 5, May 1990, pp. 473-

475. 
Table I. 65o cropped delta wing model geometry details. 

Model Geometry Details 

Wing LEX Centerline Tail Twin Tails 

Airfoil 

Modified 

NACA 64A005 

with sharp 

leading edge 

Airfoil 

Flat plate with symmetricallybeveled leading edge 

Airfoil 

Flat plate with symmetricallybeveled leading edge 

Airfoil 

Flat plate with symmetricallybeveled leading edge 

LE w, 65o LE lex , 65o LE cl , 45o LE tw, 45.24o TE w, 0o TE lex , 65o TE cl , 20o TE tw, 0o 

c 23.622 in. r lex , c 7.680 in. r cl , c 6.881 in. r tw, c 5.650 in. 

t w, c 3.544 in. t lex , c 7.680 in. t cl , c 3.175 in. t tw, c 0.880 in. 

wb 18.726 in. lex b 4.960 in. cl b 5.829 in. tw b 4.730 in. wc 16.056 in. lex c 7.680 in. cl c 5.256 in. tw c 3.846 in. 

w 0.150 lex 1.000 cl 0.461 tw 0.156 

w S 254.3553 in.2 (1.7664 ft.2) lex S 38.0928 in.2 (0.2645 ft.2) cl S 29.308 in.2 (0.2035 ft.2) tw S 15.443 in.2 (0.1072 ft.2) (per tail) 

MRC 0.57 c (M.S. 21.144) 0.25 cl c l 8.647 in. (M.S. 29.791) 0.25 tw c l 8.454 in. (M.S. 29.598) 

cham S 4.909 in.2 (0.03409 ft.2) Vcl 253.4398 in.3 (0.1467 ft.3) Vtw 261.1048 in.3 (0.1511 ft.3) 
* 8-Foot TPT Test 1057 only ** 7- by 10-Foot HST Test 202 only 

Table II. 65o cropped delta wing model pressure orifice locations. 

Pressure Orifice Locations 

M. S. 

(in.) 

x

c y (in.) y/s M. S. (in.) 

x

c y (in.) y/s M. S. (in.) 

x

c y (in.) 

14.767 

| | | | | | | | | | | | | | | | | | | | 



0.30 

| | | | | | | | | | | | | | | | | | | | 



0.000** 0.00** 21.853 

| | | | | | | | | | | | | | 

| 



0.60 

| | | | | | | | | | | | | | 

| 



0.661 0.10 26.578 

| | | | | | | | | | | | | 

 | 



0.80 

| | | | | | | | | | | | | 

| 



1.762 

0.326 0.10 1.322 0.20 2.644 

0.652* 0.20* 1.983 0.30 3.525 

0.978 0.30 2.644 0.40 4.405 1.304 0.40 3.304 0.50 4.847 1.630 0.50 3.636 0.55 5.287 1.793 0.55 3.965 0.60 5.728 

1.956** 0.60** 4.296 0.65 5.948 

2.038* 0.625* 4.626 0.70 6.167 

2.119 0.65 4.792 0.725 6.389 

2.200* 0.675* 4.957 0.75 6.830 

2.282 0.70 5.122 0.775 7.050 

2.363* 0.725* 5.287 0.80 7.490 

2.445 0.75 5.452 0.825 7.929** 2.689 0.825 5.618 0.85 8.107* 2.771 0.85 5.948 0.90 2.852 0.875 2.934 0.90 

3.000** 0.92** 3.064** 0.94** 3.130** 0.96** 
44

Table III. NASA LaRC 755 and 842A balance design loads and calibration accuracies.

M∞ q , psf ∞ CN CA Cm Cl Cn CY

0.50 282 ±0.0020 ±0.00065 ±0.00023 ±0.00019 ±0.00026 ±0.0014

0.85 430 ±0.0013 ±0.00043 ±0.00015 ±0.00013 ±0.00017 ±0.00092

1.20 521 ±0.0011 ±0.00035 ±0.00012 ±0.00010 ±0.00014 ±0.00076

Table IV. 755 balance calibration accuracies expressed in terms of aerodynamic force

 and moment coefficients.

755 Balance 

Component Design Load

(lbs or in-lbs)

Full Scale

Output

(mV)

Accuracy

% F.S.

(95% C.L.)

Accuracy

(μV)

(95% C.L.)

Accuracy

(lbs or in-lbs)

(95% C.L.)

Normal Force ±2000 12.370 0.05 6.18 1.00

Axial Force 125 4.730 0.26 12.30 0.325

Pitching Moment ±3000 6.620 0.09 5.96 2.70

Rolling Moment ±1500 5.880 0.12 7.06 1.80

Yawing Moment ±2000 9.230 0.12 11.08 2.40

Side Force ±500 6.575 0.14 9.21 0.70

842A Balance 

Component Design Load

(lbs or in-lbs)

Full Scale

Output

(mV)

Accuracy

% F.S.

(95% C.L.)

Accuracy

(μV)

(95% C.L.)

Accuracy

(lbs or in-lbs)

(95% C.L.)

Normal Force ±1600 9.350 0.09 8.42 1.44

Axial Force 75 5.920 0.40 23.68 0.30

Pitching Moment ±3000 7.255 0.10 7.26 3.00

Rolling Moment ±1500 7.490 0.10 7.49 1.50

Yawing Moment ±1500 6.630 0.17 11.27 2.55

Side Force ±500 5.395 0.19 10.25 0.95
45

M∞ ∞ q (psf) Cp uncertainty, ∆Cp

( 95% C.L.)

0.50 282 ±0.0038

0.85 430 ±0.0025

1.20 521 ±0.0021

p

p 0.0005 15 psi 144 sq.in / sq. ft. 1.08 psf C

qq q ∞∞ ∞

∆ ∗∗ ∆= = =

Table VI. Chamber pressure measurement uncertainties for 8-Foot TPT Test 1057 

 expressed in terms of the static pressure coefficient (95% confidence

 limits (C.L.) about the mean response).

M∞ Cp

* Cp,v

0.50 -2.133 -5.714

0.85 -0.302 -1.977

1.20 +0.279 -0.992

Table VII. Vacuum pressure coefficients and critical pressure coefficients 

 corresponding to the test conditions in 8-Foot TPT Test 1057.

M∞

q∞

(psf)

p∞

(psf)

po

(psf)

Re/ft

(10-6)

To

(

o

F)

0.50 162 923 1098 1.5 100

0.50 270 1539 1830 2.5 100

0.50 282 1609 1914 2.5 120

0.50 395 2252 2679 3.5 120

0.85 430 849 1364 2.5 120

1.20 521 515 1254 2.5 120

Table V. Test conditions for the 65o cropped delta wing model 

 experiment in the NASA LaRC 8-Foot TPT Test 1057.
46 

Figure 1. Several of the key components of the 65-degree cropped delta wing wind tunnel model. 

(a) Porous LEX-wing configuration with centerline tail 

Figure 2. Photographs of the 65-degree cropped delta wing model installed in the 

 NASA LaRC 8-Foot Transonic Pressure Tunnel.
47 

(b) Porous LEX-wing configuration with twin uncanted tails 

Figure 2. Continued. 

(c) Porous LEX-wing configuration with twin canted tails 

Figure 2. Continued. 
48 

(d) Wing-alone configuration with centerline tail 

Figure 2. Continued. 

(e) Canard-wing configuration with centerline tail

Figure 2. Continued. 
49 

(f) Close-up photographs of LEX lower surface with gooseneck attachment to wing (left)

and wing lower surface with LEX removed (right)

Figure 2. Continued. 

(g) Miniature video camera and housing installed on model support system 

Figure 2. Concluded. 
50 

Figure 3. Planview of the 65o cropped delta wing model with LEX. (Dimensions are in inches.) 

15 psid 48-port

ESP module

M.S.

0.00

M.S.

21.144

M.S.

31.302

M.S.

32.286

No. 20 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.

21.811

755 BMC

W.L.

4.131

Figure 4. Sideview of the 65o cropped delta wing model with centerline vertical tail. 

 (Dimensions are in inches.) 

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.

9.363

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°

SAREA1=254.3553 sq.in.

CHORD1=23.622 in.

BSPAN1=18.726 in.

XBAR1=0.667 in.

YBAR1=0.00 in.

ZBAR1=0.984 in.

ARPBA1.1=4.9090 sq.in.
51 

M.S.

0.00

M.S.

21.144

M.S.

31.302

M.S.

32.286

755 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

Figure 5. Sideview of the 65o

 delta wing model with twin vertical tails. 

 (Dimensions are in inches.) 

Moment

Reference

Center (MRC)

Centerline

Vertical Tail

A A

Section A-A

M.S.

0.593

M.S.

7.680

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.

9.363

M.S.

21.144

B.L.

4.091

Canard

65°

90°

90°

M.S.

10.545

60°

35°

t/c = 0.05 

Figure 6. Planview of the 65o

 cropped delta wing model with canard. (Dimensions are in inches.) 
52 

M.S.

0.00

M.S.

5.318

M.S.

12.998

(a) porosity level 2 

M.S.

0.00

M.S.

5.318

M.S.

12.998

(b) porosity level 3 

Figure 7. Compartmentalized porosity on LEX. 
53 

M.S.

0.00

M.S.

5.318

M.S.

12.998

0.575

0.575

(c) porosity level 4 

M.S.

0.00

M.S.

5.318

M.S.

12.998

0.575

0.575

(d) porosity level 5 

 Figure 7. Concluded. 
54 

26.578

x/c=0.80

65°

65°

90°

90°

y

x

M.S. 

14.767

x/c=0.30

M.S. 

21.853

x/c=0.60

Figure 8. Pressure measurement stations on the 65o cropped delta wing model. 

3.305

6.609

8.812

x/c=0.30

M.S. 21.853

x/c=0.60

M.S. 26.578

x/c=0.80

W.L. 0.000

W.L. 0.984

W.L. 0.984

W.L. 0.000

W.L. 0.984

W.L. 0.000

y

y

y

Fuselage

Figure 9. Cross sections of the 65o

 cropped delta wing model at the three pressure measurement 

 stations. (Dimensions are in inches.) 
55 

12.661 2.000

45.00°

45.00°

pcham1

Balance

Cable

pcham2

ESP & TC wiring and tubing

(both sides) -- use metal tape or

nichrome strips to secure to

fuselage inner surface

Fouling Strip

Balance Dowel Hole

12.161

Chamber

Tubes (2)

External

to Sting

Balance

Cable

(Internal

to Sting)

Figure 10. Chamber pressure tube and fouling strip layout on the 65o cropped delta wing model. 

 (Dimensions are in inches.) 

(a) high angle-of-attack static stability support system 

Figure 11. Photographs of the 65o cropped delta wing model installed in the 

 NASA LaRC 7- by 10-Foot High Speed Tunnel. 
56 

(b) 65o cropped delta wing model with porous LEX and twin canted tails 

Figure 11. Continued. 

(c) 65o

 cropped delta wing model with twin canted tails; LEX off 

Figure 11. Concluded. 
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid 

Re/ft (10-6) 

1.5 2.5 3.5 

α, deg 

10.00 

9.95 

10.11 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 10 degrees angle of attack Figure 12. Effect of the Reynolds number on the wing upper surface static pressure distributions at Mach = 0.50 with solid 
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid 

Re/ft (10-6) 

1.5 2.5 3.5 

α, deg 

19.98 20.10 20.05 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(b) 20 degrees angle of attack  Figure 12. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous 

Re/ft (10-6) 

1.5 2.5 3.5 

α, deg 

10.02 10.07 10.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 10 degrees angle of attack Figure 13. Effect of the Reynolds number on the wing upper surface static pressure distributions at Mach = 0.50 with porou
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous 

Re/ft (10-6) 

1.5 2.5 3.5 

α, deg 

20.17 20.05 20.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(b) 20 degrees angle of attack  Figure 13. Concluded.
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10 

10 

10 

10 

10 

10 

Point 

400 

400 

400 

369 

369 

369 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

1.5 

1.5 

1.5 

3.5 

3.5 

3.5 

α, deg 

10.11 

10.11 

10.11 

10.00 

10.00 

10.00 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 21:40:33 2010 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

.16 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 (a) 10 degrees angle of attack 

Figure 14. Upper surface static pressure coefficient comparisons at Re/ft = 1.5 million and 3.5 million 

 and Mach = 0.50 with solid LEX and centerline tail.

 61
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10 

10 

10 

10 

10 

10 

Point 

405 

405 

405 

374 

374 

374 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

1.5 

1.5 

1.5 

3.5 

3.5 

3.5 

α, deg 

20.05 

20.05 

20.05 

19.98 

19.98 

19.98 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 21:31:31 2010 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

.16 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

(b) 20 degrees angle of attack 

 Figure 14. Concluded.

 62
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10 

10 

10 

10 

10 

10 

Point 

400 

400 

400 

384 

384 

384 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.5 

3.5 

3.5 

α, deg 

10.11 

10.11 

10.11 

9.95 

9.95 

9.95 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Sat Jul 17 15:07:04 2010 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

.16 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 (a) 10 degrees angle of attack 

Figure 15. Upper surface static pressure coefficient comparisons at Re/ft = 2.5 million and 3.5 million 

 and Mach = 0.50 with solid LEX and centerline tail.

 63
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10 

10 

10 

10 

10 

10 

Point 

405 

405 

405 

389 

389 

389 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.5 

3.5 

3.5 

α, deg 

20.05 

20.05 

20.05 

20.10 

20.10 

20.10 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Sat Jul 17 15:04:41 2010 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

.16 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

(b) 20 degrees angle of attack 

 Figure 15. Concluded.

 64
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

4 

4 

4 

3 

3 

3 

Point 

73 

73 

73 

26 

26 

26 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

1.5 

1.5 

1.5 

3.5 

3.5 

3.5 

α, deg 

10.03 

10.03 

10.03 

10.02 

10.02 

10.02 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 21:49:12 2010 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

.16 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 (a) 10 degrees angle of attack 

Figure 16. Upper surface static pressure coefficient comparisons at Re/ft = 1.5 million and 3.5 million 

 and Mach = 0.50 with porous LEX and centerline tail.

 65
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

4 

4 

4 

3 

3 

3 

Point 

78 

78 

78 

31 

31 

31 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

1.5 

1.5 

1.5 

3.5 

3.5 

3.5 

α, deg 

20.04 

20.04 

20.04 

20.17 

20.17 

20.17 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 21:46:01 2010 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

.16 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

(b) 20 degrees angle of attack 

 Figure 16. Concluded.

 66
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

4 

4 

4 

3 

3 

3 

Point 

73 

73 

73 

40 

40 

40 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.5 

3.5 

3.5 

α, deg 

10.03 

10.03 

10.03 

10.07 

10.07 

10.07 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Sat Jul 17 15:01:23 2010 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

.16 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 (a) 10 degrees angle of attack 

Figure 17. Upper surface static pressure coefficient comparisons at Re/ft = 2.5 million and 3.5 million 

 and Mach = 0.50 with porous LEX and centerline tail.

 67
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

4 

4 

4 

3 

3 

3 

Point 

78 

78 

78 

45 

45 

45 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.5 

3.5 

3.5 

α, deg 

20.04 

20.04 

20.04 

20.05 

20.05 

20.05 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Sat Jul 17 14:55:57 2010 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

.16 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

(b) 20 degrees angle of attack 

 Figure 17. Concluded.

 68
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

LEX 

Solid 

Solid 

Solid 

Re/ft (10-6) 

1.5 

2.5 

3.5 

Figure 18. Effect of the Reynolds number on the longitudinal aerodynamic characteristics at Mach = 0.50 

 with solid LEX and centerline tail. 69
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

LEX 

Porous 

Porous 

Porous 

Re/ft (10-6) 

1.5 

2.5 

3.5 

Figure 19. Effect of the Reynolds number on the longitudinal aerodynamic characteristics at Mach = 0.50 

 with porous LEX and centerline tail. 70
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.8 

-.4 

0 

.4 

.8 

1.2 

1.6 

2.0 

CN 

Test No. 

1057. 

1057. 

1057. 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10. 

10. 

10. 

Polar 

22. 

23. 

24. 

M∞ 

0.50 

0.50 

0.50 

LEX 

Solid 

Solid 

Solid 

Re/ft (10-6) 

1.5 

2.5 

3.5 

 Tests 1057 Porous LEX Model NASA LaRC 8-Foot TPT 

Mon Jul 19 19:41:48 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.02 

-.01 

0 

.01 

.02 

ΔCN 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (a) normal force coefficient 

Figure 20. Comparison of the delta longitudinal aerodynamic coefficients at Re/ft = 1.5 million, 

 2.5 million, and 3.5 million and Mach = 0.50 with solid LEX and centerline tail.

 71
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

0 

.004 

.008 

.012 

.016 

.020 

.024 

.028 

CA 

Test No. 

1057. 

1057. 

1057. 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10. 

10. 

10. 

Polar 

22. 

23. 

24. 

M∞ 

0.50 

0.50 

0.50 

LEX 

Solid 

Solid 

Solid 

Re/ft (10-6) 

1.5 

2.5 

3.5 

 Tests 1057 Porous LEX Model NASA LaRC 8-Foot TPT 

Mon Jul 19 19:45:41 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCA 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

(b) axial force coefficient 

 Figure 20. Continued.

 72
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

0 

.02 

.04 

.06 

.08 

.10 

.12 

.14 

Cm 

Test No. 

1057. 

1057. 

1057. 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10. 

10. 

10. 

Polar 

22. 

23. 

24. 

M∞ 

0.50 

0.50 

0.50 

LEX 

Solid 

Solid 

Solid 

Re/ft (10-6) 

1.5 

2.5 

3.5 

 Tests 1057 Porous LEX Model NASA LaRC 8-Foot TPT 

Mon Jul 19 21:03:37 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCm 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (c) pitching moment coefficient 

 Figure 20. Concluded.

 73
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.8 

-.4 

0 

.4 

.8 

1.2 

1.6 

2.0 

CN 

Test No. 

1057. 

1057. 

1057. 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

3. 

3. 

4. 

Polar 

1. 

2. 

4. 

M∞ 

0.50 

0.50 

0.50 

LEX 

Porous 

Porous 

Porous 

Re/ft (10-6) 

1.5 

2.5 

3.5 

 Tests 1057 Porous LEX Model NASA LaRC 8-Foot TPT 

Mon Jul 19 19:16:26 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.02 

-.01 

0 

.01 

.02 

ΔCN 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (a) normal force coefficient 

Figure 21. Comparison of the delta longitudinal aerodynamic coefficients at Re/ft = 1.5 million, 

 2.5 million, and 3.5 million and Mach = 0.50 with porous LEX and centerline tail.

 74
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

.012 

.016 

.020 

.024 

.028 

.032 

.036 

.040 

CA 

Test No. 

1057. 

1057. 

1057. 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

3. 

3. 

4. 

Polar 

1. 

2. 

4. 

M∞ 

0.50 

0.50 

0.50 

LEX 

Porous 

Porous 

Porous 

Re/ft (10-6) 

1.5 

2.5 

3.5 

 Tests 1057 Porous LEX Model NASA LaRC 8-Foot TPT 

Mon Jul 19 19:36:15 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCA 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

(b) axial force coefficient 

 Figure 21. Continued.

 75
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

0 

.02 

.04 

.06 

.08 

.10 

.12 

.14 

Cm 

Test No. 

1057. 

1057. 

1057. 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

3. 

3. 

4. 

Polar 

1. 

2. 

4. 

M∞ 

0.50 

0.50 

0.50 

LEX 

Porous 

Porous 

Porous 

Re/ft (10-6) 

1.5 

2.5 

3.5 

 Tests 1057 Porous LEX Model NASA LaRC 8-Foot TPT 

Mon Jul 19 21:07:18 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCm 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (c) pitching moment coefficient 

 Figure 21. Concluded.

 76
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

10. 11. 

Polar 

23. 36. 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 2.5 

α, deg 

12.06 12.01 

LEX 

Solid Solid 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 12 degrees angle of attack Figure 22. Repeatability of the wing upper surface static pressure distributions at Mach = 0.50 and Re/ft = 2.5 million with so
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

10. 11. 

Polar 

23. 36. 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 2.5 

α, deg 

16.08 16.04 

LEX 

Solid Solid 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(b) 16 degrees angle of attack  Figure 22. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

10. 11. 

Polar 

23. 36. 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 2.5 

α, deg 

20.10 20.03 

LEX 

Solid Solid 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(c) 20 degrees angle of attack  Figure 22. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

10. 11. 

Polar 

23. 36. 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 2.5 

α, deg 

24.10 24.06 

LEX 

Solid Solid 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(d) 24 degrees angle of attack  Figure 22. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

10. 11. 

Polar 

21. 30. 

M∞ 

0.85 0.85 

Re/ft (10-6) 

2.5 2.5 

α, deg 

12.10 12.02 

LEX 

Solid Solid 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 12 degrees angle of attack Figure 23. Repeatability of the wing upper surface static pressure distributions at Mach = 0.85 and Re/ft = 2.5 million with so
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

10. 11. 

Polar 

21. 30. 

M∞ 

0.85 0.85 

Re/ft (10-6) 

2.5 2.5 

α, deg 

16.10 16.12 

LEX 

Solid Solid 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(b) 16 degrees angle of attack  Figure 23. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

10. 11. 

Polar 

21. 30. 

M∞ 

0.85 0.85 

Re/ft (10-6) 

2.5 2.5 

α, deg 

20.10 20.13 

LEX 

Solid Solid 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(c) 20 degrees angle of attack  Figure 23. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

10. 11. 

Polar 

21. 30. 

M∞ 

0.85 0.85 

Re/ft (10-6) 

2.5 2.5 

α, deg 

23.98 24.09 

LEX 

Solid Solid 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(d) 24 degrees angle of attack  Figure 23. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

3. 7. 

Polar 

2. 

16. 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 2.5 

α, deg 

12.07 12.01 

LEX 

PorouPorou

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 12 degrees angle of attack Figure 24. Repeatability of the wing upper surface static pressure distributions at Mach = 0.50 and Re/ft = 2.5 million with por
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

3. 7. 

Polar 

2. 

16. 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 2.5 

α, deg 

16.05 16.08 

LEX 

PorouPorou

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(b) 16 degrees angle of attack  Figure 24. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

3. 7. 

Polar 

2. 

16. 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 2.5 

α, deg 

20.05 20.05 

LEX 

PorouPorou

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(c) 20 degrees angle of attack  Figure 24. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

3. 7. 

Polar 

2. 

16. 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 2.5 

α, deg 

24.03 24.04 

LEX 

PorouPorou

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(d) 24 degrees angle of attack  Figure 24. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

4. 6. 

Polar 

3. 

11. 

M∞ 

0.85 0.85 

Re/ft (10-6) 

2.5 2.5 

α, deg 

12.02 12.05 

LEX 

PorouPorou

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 12 degrees angle of attack Figure 25. Repeatability of the wing upper surface static pressure distributions at Mach = 0.85 and Re/ft = 2.5 million with por
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

4. 6. 

Polar 

3. 

11. 

M∞ 

0.85 0.85 

Re/ft (10-6) 

2.5 2.5 

α, deg 

16.04 16.00 

LEX 

PorouPorou

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(b) 16 degrees angle of attack  Figure 25. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

4. 6. 

Polar 

3. 

11. 

M∞ 

0.85 0.85 

Re/ft (10-6) 

2.5 2.5 

α, deg 

20.07 20.03 

LEX 

PorouPorou

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(c) 20 degrees angle of attack  Figure 25. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 1057. 

Run 

4. 6. 

Polar 

3. 

11. 

M∞ 

0.85 0.85 

Re/ft (10-6) 

2.5 2.5 

α, deg 

24.06 24.02 

LEX 

PorouPorou

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(d) 24 degrees angle of attack  Figure 25. Concluded.
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10 

10 

10 

11 

11 

11 

Point 

385 

385 

385 

615 

615 

615 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

12.06 

12.06 

12.06 

12.01 

12.01 

12.01 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 15:02:23 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 (a) 12 degrees angle of attack 

Figure 26. Delta pressure coefficient distribution in repeat runs at Mach = 0.50 and Re/ft = 2.5 million 

 with solid LEX and centerline tail.

 93
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10 

10 

10 

11 

11 

11 

Point 

387 

387 

387 

617 

617 

617 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

16.08 

16.08 

16.08 

16.04 

16.04 

16.04 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 15:01:04 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

(b) 16 degrees angle of attack 

 Figure 26. Continued.

 94
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10 

10 

10 

11 

11 

11 

Point 

389 

389 

389 

619 

619 

619 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

20.10 

20.10 

20.10 

20.03 

20.03 

20.03 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 14:57:23 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

(c) 20 degrees angle of attack 

 Figure 26. Continued.

 95
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10 

10 

10 

11 

11 

11 

Point 

391 

391 

391 

621 

621 

621 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

24.10 

24.10 

24.10 

24.06 

24.06 

24.06 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 14:46:55 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

(d) 24 degrees angle of attack 

 Figure 26. Concluded.

 96
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10 

10 

10 

11 

11 

11 

Point 

355 

355 

355 

509 

509 

509 

M∞ 

0.85 

0.85 

0.85 

0.85 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

12.10 

12.10 

12.10 

12.02 

12.02 

12.02 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 18:33:51 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0025 at M = 0.85 

 (a) 12 degrees angle of attack 

Figure 27. Delta pressure coefficient distribution in repeat runs at Mach = 0.85 and Re/ft = 2.5 million 

 with solid LEX and centerline tail.

 97
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10 

10 

10 

11 

11 

11 

Point 

357 

357 

357 

511 

511 

511 

M∞ 

0.85 

0.85 

0.85 

0.85 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

16.10 

16.10 

16.10 

16.12 

16.12 

16.12 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 17:41:18 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0025 at M = 0.85 

(b) 16 degrees angle of attack 

 Figure 27. Continued.

 98
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10 

10 

10 

11 

11 

11 

Point 

359 

359 

359 

513 

513 

513 

M∞ 

0.85 

0.85 

0.85 

0.85 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

20.10 

20.10 

20.10 

20.13 

20.13 

20.13 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 17:38:32 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0025 at M = 0.85 

(c) 20 degrees angle of attack 

 Figure 27. Continued.

 99
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

10 

10 

10 

11 

11 

11 

Point 

361 

361 

361 

515 

515 

515 

M∞ 

0.85 

0.85 

0.85 

0.85 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

23.98 

23.98 

23.98 

24.09 

24.09 

24.09 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 17:26:36 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0025 at M = 0.85 

(d) 24 degrees angle of attack 

 Figure 27. Concluded.

 100
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

3 

3 

3 

7 

7 

7 

Point 

41 

41 

41 

269 

269 

269 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

12.07 

12.07 

12.07 

12.01 

12.01 

12.01 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 18:52:32 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 (a) 12 degrees angle of attack 

Figure 28. Delta pressure coefficient distribution in repeat runs at Mach = 0.50 and Re/ft = 2.5 million 

 with porous LEX and centerline tail.

 101
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

3 

3 

3 

7 

7 

7 

Point 

43 

43 

43 

271 

271 

271 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

16.05 

16.05 

16.05 

16.08 

16.08 

16.08 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 18:48:14 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

(b) 16 degrees angle of attack 

 Figure 28. Continued.

 102
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

3 

3 

3 

7 

7 

7 

Point 

45 

45 

45 

273 

273 

273 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

20.05 

20.05 

20.05 

20.05 

20.05 

20.05 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 18:45:33 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

(c) 20 degrees angle of attack 

 Figure 28. Continued.

 103
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

3 

3 

3 

7 

7 

7 

Point 

47 

47 

47 

275 

275 

275 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

24.03 

24.03 

24.03 

24.04 

24.04 

24.04 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 18:42:49 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

(d) 24 degrees angle of attack 

 Figure 28. Concluded.

 104
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

4 

4 

4 

6 

6 

6 

Point 

60 

60 

60 

187 

187 

187 

M∞ 

0.85 

0.85 

0.85 

0.85 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

12.02 

12.02 

12.02 

12.05 

12.05 

12.05 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 19:03:28 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0025 at M = 0.85 

 (a) 12 degrees angle of attack 

Figure 29. Delta pressure coefficient distribution in repeat runs at Mach = 0.85 and Re/ft = 2.5 million 

 with porous LEX and centerline tail.

 105
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

4 

4 

4 

6 

6 

6 

Point 

62 

62 

62 

189 

189 

189 

M∞ 

0.85 

0.85 

0.85 

0.85 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

16.04 

16.04 

16.04 

16.00 

16.00 

16.00 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 19:01:31 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0025 at M = 0.85 

(b) 16 degrees angle of attack 

 Figure 29. Continued.

 106
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

4 

4 

4 

6 

6 

6 

Point 

64 

64 

64 

191 

191 

191 

M∞ 

0.85 

0.85 

0.85 

0.85 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

20.07 

20.07 

20.07 

20.03 

20.03 

20.03 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 18:58:53 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0025 at M = 0.85 

(c) 20 degrees angle of attack 

 Figure 29. Continued.

 107
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

1057 

1057 

1057 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

Run 

4 

4 

4 

6 

6 

6 

Point 

66 

66 

66 

193 

193 

193 

M∞ 

0.85 

0.85 

0.85 

0.85 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

2.5 

2.5 

2.5 

2.5 

α, deg 

24.06 

24.06 

24.06 

24.02 

24.02 

24.02 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

plot_dq1057b 

Fri Jul 16 18:55:12 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0025 at M = 0.85 

(d) 24 degrees angle of attack 

 Figure 29. Concluded.

 108
0 4 8 12 16 20 24 28 32 

α, deg 

-.8 

-.4 

0 

.4 

.8 

1.2 

1.6 

2.0 

CN 

Test No. 

1057. 

1057. 

Run 

10. 

11. 

Polar 

23. 

36. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

LEX 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

 Test 1057 Porous LEX Model NASA LaRC 8-Foot Transonic Pressure Tunnel 

Mon Jul 12 12:49:56 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.010 

-.005 

0 

.005 

.010 

ΔCN 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (a) normal force coefficient 

Figure 30. Comparison of the delta longitudinal aerodynamic coefficients in repeat runs at Mach = 0.50 

 and Re/ft = 2.5 million with solid LEX and centerline tail.

 109
0 4 8 12 16 20 24 28 32 

α, deg 

0 

.002 

.004 

.006 

.008 

.010 

.012 

.014 

CA 

Test No. 

1057. 

1057. 

Run 

10. 

11. 

Polar 

23. 

36. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

LEX 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

 Test 1057 Porous LEX Model NASA LaRC 8-Foot Transonic Pressure Tunnel 

Mon Jul 12 10:34:41 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCA 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

(b) axial force coefficient 

 Figure 30. Continued.

 110
0 4 8 12 16 20 24 28 32 

α, deg 

0 

.02 

.04 

.06 

.08 

.10 

.12 

.14 

Cm 

Test No. 

1057. 

1057. 

Run 

10. 

11. 

Polar 

23. 

36. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

LEX 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

 Test 1057 Porous LEX Model NASA LaRC 8-Foot Transonic Pressure Tunnel 

Mon Jul 12 10:03:06 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCm 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (c) pitching moment coefficient 

 Figure 30. Concluded.

 111
0 4 8 12 16 20 24 28 32 

α, deg 

-.8 

-.4 

0 

.4 

.8 

1.2 

1.6 

2.0 

CN 

Test No. 

1057. 

1057. 

Run 

10. 

11. 

Polar 

21. 

30. 

M∞ 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

LEX 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

 Test 1057 Porous LEX Model NASA LaRC 8-Foot Transonic Pressure Tunnel 

Mon Jul 12 13:16:40 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.010 

-.005 

0 

.005 

.010 

ΔCN 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (a) normal force coefficient 

Figure 31. Comparison of the delta longitudinal aerodynamic coefficients in repeat runs at Mach = 0.85 

 and Re/ft = 2.5 million with solid LEX and centerline tail.

 112
0 4 8 12 16 20 24 28 32 

α, deg 

0 

.004 

.008 

.012 

.016 

.020 

.024 

.028 

CA 

Test No. 

1057. 

1057. 

Run 

10. 

11. 

Polar 

21. 

30. 

M∞ 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

LEX 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

 Test 1057 Porous LEX Model NASA LaRC 8-Foot Transonic Pressure Tunnel 

Mon Jul 12 10:31:09 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCA 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

(b) axial force coefficient 

 Figure 31. Continued.

 113
0 4 8 12 16 20 24 28 32 

α, deg 

0 

.02 

.04 

.06 

.08 

.10 

.12 

.14 

Cm 

Test No. 

1057. 

1057. 

Run 

10. 

11. 

Polar 

21. 

30. 

M∞ 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

LEX 

Solid 

Solid 

Vertical Tail 

Centerline Tail 

Centerline Tail 

 Test 1057 Porous LEX Model NASA LaRC 8-Foot Transonic Pressure Tunnel 

Mon Jul 12 12:30:18 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCm 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (c) pitching moment coefficient 

 Figure 31. Concluded.

 114
0 4 8 12 16 20 24 28 32 

α, deg 

-.8 

-.4 

0 

.4 

.8 

1.2 

1.6 

2.0 

CN 

Test No. 

1057. 

1057. 

Run 

3. 

7. 

Polar 

2. 

16. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

LEX 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

 Test 1057 Porous LEX Model NASA LaRC 8-Foot Transonic Pressure Tunnel 

Mon Jul 12 13:30:55 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.010 

-.005 

0 

.005 

.010 

ΔCN 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (a) normal force coefficient 

Figure 32. Comparison of the delta longitudinal aerodynamic coefficients in repeat runs at Mach = 0.50 

 and Re/ft = 2.5 million with porous LEX and centerline tail.

 115
0 4 8 12 16 20 24 28 32 

α, deg 

.012 

.016 

.020 

.024 

.028 

.032 

.036 

.040 

CA 

Test No. 

1057. 

1057. 

Run 

3. 

7. 

Polar 

2. 

16. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

LEX 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

 Test 1057 Porous LEX Model NASA LaRC 8-Foot Transonic Pressure Tunnel 

Mon Jul 12 13:41:14 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCA 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

(b) axial force coefficient 

 Figure 32. Continued.

 116
0 4 8 12 16 20 24 28 32 

α, deg 

0 

.02 

.04 

.06 

.08 

.10 

.12 

.14 

Cm 

Test No. 

1057. 

1057. 

Run 

3. 

7. 

Polar 

2. 

16. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

LEX 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

 Test 1057 Porous LEX Model NASA LaRC 8-Foot Transonic Pressure Tunnel 

Mon Jul 12 13:42:54 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCm 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (c) pitching moment coefficient 

 Figure 32. Concluded.

 117
0 4 8 12 16 20 24 28 32 

α, deg 

-.8 

-.4 

0 

.4 

.8 

1.2 

1.6 

2.0 

CN 

Test No. 

1057. 

1057. 

Run 

4. 

6. 

Polar 

3. 

11. 

M∞ 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

LEX 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

 Test 1057 Porous LEX Model NASA LaRC 8-Foot Transonic Pressure Tunnel 

Mon Jul 12 14:32:09 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.010 

-.005 

0 

.005 

.010 

ΔCN 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (a) normal force coefficient 

Figure 33. Comparison of the delta longitudinal aerodynamic coefficients in repeat runs at Mach = 0.85 

 and Re/ft = 2.5 million with porous LEX and centerline tail.

 118
0 4 8 12 16 20 24 28 32 

α, deg 

.012 

.016 

.020 

.024 

.028 

.032 

.036 

.040 

CA 

Test No. 

1057. 

1057. 

Run 

4. 

6. 

Polar 

3. 

11. 

M∞ 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

LEX 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

 Test 1057 Porous LEX Model NASA LaRC 8-Foot Transonic Pressure Tunnel 

Tue Aug 24 16:07:09 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCA 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

(b) axial force coefficient 

 Figure 33. Continued.

 119
0 4 8 12 16 20 24 28 32 

α, deg 

0 

.004 

.008 

.012 

.016 

.020 

.024 

.028 

Cm 

Test No. 

1057. 

1057. 

Run 

4. 

6. 

Polar 

3. 

11. 

M∞ 

0.85 

0.85 

Re/ft (10-6) 

2.5 

2.5 

LEX 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

 Test 1057 Porous LEX Model NASA LaRC 8-Foot Transonic Pressure Tunnel 

Mon Jul 12 15:06:00 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCm 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (c) pitching moment coefficient 

 Figure 33. Concluded.

 120
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

10.02 10.13 

LEX 

Solid Solid 

Ve

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 10 degrees angle of attack Figure 34. Tunnel-to-tunnel comparison of the wing upper surface static pressure distributions at Mach = 0.50 with s
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

20.03 20.04 

LEX 

Solid Solid 

V

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 20 degrees angle of attack  Figure 34. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

10.04 10.04 

LEX 

Porous Porous 

V

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 10 degrees angle of attack Figure 35. Tunnel-to-tunnel comparison of the wing upper surface static pressure distributions at Mach = 0.50 wit centerline tail on (T1057), tail off (T202).
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

20.05 20.07 

LEX 

Porous Porous 

V

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 20 degrees angle of attack  Figure 35. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

10.00 10.08 

LE

OfOf

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (a) 10 degrees angle of attack Figure 36. Tunnel-to-tunnel comparison of the wing upper surface static pressure distributions at Mach = 0.50 wi centerline tail on (T1057), tail off (T202).
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

20.03 20.07 

LE

OfOf

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 20 degrees angle of attack  Figure 36. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

10.02 10.09 

LE

SolSol

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 10 degrees angle of attack Figure 37. Tunnel-to-tunnel comparison of the wing upper surface static pressure distributions at Mach = 0.50 with solid LEX 
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

20.03 20.10 

LE

SolSol

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 20 degrees angle of attack  Figure 37. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

10.05 10.03 

LE

SolSol

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 10 degrees angle of attack Figure 38. Tunnel-to-tunnel comparison of the wing upper surface static pressure distributions at Mach = 0.50 with solid LE
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

20.06 20.19 

LE

SolSol

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 20 degrees angle of attack  Figure 38. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

10.08 10.15 

LE

PorPor

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 10 degrees angle of attack Figure 39. Tunnel-to-tunnel comparison of the wing upper surface static pressure distributions at Mach = 0.50 with porous LE
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

20.02 20.07 

LE

PorPor

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 20 degrees angle of attack  Figure 39. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

10.02 10.16 

LE

OfOf

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (a) 10 degrees angle of attack Figure 40. Tunnel-to-tunnel comparison of the wing upper surface static pressure distributions at Mach = 0.50 with wing alo
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

M∞ 

0.50 0.50 

Re/ft (10-6) 

2.5 3.0 

α, deg 

20.02 20.16 

LE

OfOf

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 20 degrees angle of attack  Figure 40. Concluded.
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

23 

23 

23 

8 

8 

8 

Point 

1377 

1377 

1377 

277 

277 

277 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

10.02 

10.02 

10.02 

10.13 

10.13 

10.13 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Tail Off 

Tail Off 

Tail Off 

Tail Off 

Tail Off 

Tail Off 

plot_dq1057d 

Thu Jul 15 18:53:32 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

 (a) 10 degrees angle of attack 

Figure 41. Tunnel-to-tunnel delta pressure coefficients at Mach = 0.50 with solid LEX; tail off.

 135
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

23 

23 

23 

8 

8 

8 

Point 

1383 

1383 

1383 

281 

281 

281 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

20.03 

20.03 

20.03 

20.04 

20.04 

20.04 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Tail Off 

Tail Off 

Tail Off 

Tail Off 

Tail Off 

Tail Off 

plot_dq1057d 

Thu Jul 15 18:50:20 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

(b) 20 degrees angle of attack 

 Figure 41. Concluded.

 136
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

7 

7 

7 

12 

12 

12 

Point 

268 

268 

268 

580 

580 

580 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

10.04 

10.04 

10.04 

10.04 

10.04 

10.04 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Tail Off 

Tail Off 

Tail Off 

plot_dq1057d 

Thu Jul 15 20:12:57 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

 (a) 10 degrees angle of attack 

Figure 42. Tunnel-to-tunnel delta pressure coefficients at Mach = 0.50 with porous LEX; 

 centerline tail on (T1057), tail off (T202).

 137
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

7 

7 

7 

12 

12 

12 

Point 

273 

273 

273 

584 

584 

584 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

20.05 

20.05 

20.05 

20.07 

20.07 

20.07 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Tail Off 

Tail Off 

Tail Off 

plot_dq1057d 

Thu Jul 15 20:08:26 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

(b) 20 degrees angle of attack 

 Figure 42. Concluded.

 138
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

33 

33 

33 

23 

23 

23 

Point 

2066 

2066 

2066 

1015 

1015 

1015 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

10.00 

10.00 

10.00 

10.08 

10.08 

10.08 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Off 

Off 

Off 

Off 

Off 

Off 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Tail Off 

Tail Off 

Tail Off 

plot_dq1057d 

Thu Jul 15 21:06:54 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

 (a) 10 degrees angle of attack 

Figure 43. Tunnel-to-tunnel delta pressure coefficients at Mach = 0.50 with wing alone; 

 centerline tail on (T1057), tail off (T202).

 139
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.8 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

33 

33 

33 

23 

23 

23 

Point 

2071 

2071 

2071 

1019 

1019 

1019 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

20.03 

20.03 

20.03 

20.07 

20.07 

20.07 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Off 

Off 

Off 

Off 

Off 

Off 

Vertical Tail 

Centerline Tail 

Centerline Tail 

Centerline Tail 

Tail Off 

Tail Off 

Tail Off 

plot_dq1057d 

Thu Jul 15 21:03:31 2010 

-.40 

-.32 

-.24 

-.16 

-.08 

0 

.08 

.16 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

(b) 20 degrees angle of attack 

 Figure 43. Concluded.

 140
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

20 

20 

20 

15 

15 

15 

Point 

1170 

1170 

1170 

659 

659 

659 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

10.02 

10.02 

10.02 

10.09 

10.09 

10.09 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Twin δv=0o

Twin δv=0o

Twin δv=0o

Twin δv=0o

Twin δv=0o

Twin δv=0o

plot_dq1057d 

Thu Jul 15 19:03:10 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

 (a) 10 degrees angle of attack 

Figure 44. Tunnel-to-tunnel delta pressure coefficients at Mach = 0.50 with solid LEX and twin uncanted tails.

 141
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

20 

20 

20 

15 

15 

15 

Point 

1175 

1175 

1175 

663 

663 

663 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

20.03 

20.03 

20.03 

20.10 

20.10 

20.10 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Twin δv=0o

Twin δv=0o

Twin δv=0o

Twin δv=0o

Twin δv=0o

Twin δv=0o

plot_dq1057d 

Thu Jul 15 18:59:13 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

(b) 20 degrees angle of attack 

 Figure 44. Concluded.

 142
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

26 

26 

26 

5 

5 

5 

Point 

1573 

1573 

1573 

154 

154 

154 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

10.05 

10.05 

10.05 

10.03 

10.03 

10.03 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Twin δv=30o

Twin δv=30o

Twin δv=30o

Twin δv=30o

Twin δv=30o

Twin δv=30o

plot_dq1057d 

Thu Jul 15 16:51:29 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

 (a) 10 degrees angle of attack 

Figure 45. Tunnel-to-tunnel delta pressure coefficients at Mach = 0.50 with solid LEX and twin canted tails.

 143
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

26 

26 

26 

5 

5 

5 

Point 

1578 

1578 

1578 

158 

158 

158 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

20.06 

20.06 

20.06 

20.19 

20.19 

20.19 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Twin δv=30o

Twin δv=30o

Twin δv=30o

Twin δv=30o

Twin δv=30o

Twin δv=30o

plot_dq1057d 

Thu Jul 15 16:54:25 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

(b) 20 degrees angle of attack 

 Figure 45. Concluded.

 144
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

17 

17 

17 

14 

14 

14 

Point 

811 

811 

811 

613 

613 

613 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

10.08 

10.08 

10.08 

10.15 

10.15 

10.15 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Twin δv=0o

Twin δv=0o

Twin δv=0o

Twin δv=0o

Twin δv=0o

Twin δv=0o

plot_dq1057d 

Thu Jul 15 19:52:00 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

 (a) 10 degrees angle of attack 

Figure 46. Tunnel-to-tunnel delta pressure coefficients at Mach = 0.50 with porous LEX and twin uncanted tails.

 145
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

17 

17 

17 

14 

14 

14 

Point 

816 

816 

816 

617 

617 

617 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

20.02 

20.02 

20.02 

20.07 

20.07 

20.07 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Twin δv=0o

Twin δv=0o

Twin δv=0o

Twin δv=0o

Twin δv=0o

Twin δv=0o

plot_dq1057d 

Thu Jul 15 19:49:05 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

(b) 20 degrees angle of attack 

 Figure 46. Concluded.

 146
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

28 

28 

28 

25 

25 

25 

Point 

1803 

1803 

1803 

1106 

1106 

1106 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

10.02 

10.02 

10.02 

10.16 

10.16 

10.16 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Off 

Off 

Off 

Off 

Off 

Off 

Vertical Tail 

Twin δv=30o

Twin δv=30o

Twin δv=30o

Twin δv=30o

Twin δv=30o

Twin δv=30o

plot_dq1057d 

Thu Jul 15 21:56:55 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

 (a) 10 degrees angle of attack 

Figure 47. Tunnel-to-tunnel delta pressure coefficients at Mach = 0.50 with wing alone and twin canted tails.

 147
0 5 10 15 20 25 30 35 40 45 50 

Orifice Number 

-2.4 

-2.0 

-1.6 

-1.2 

-.8 

-.4 

0 

.4 

Cp 

Test No. 

1057 

1057 

1057 

202 

202 

202 

Facility 

8-Foot TPT 

8-Foot TPT 

8-Foot TPT 

7x10-Foot HST 

7x10-Foot HST 

7x10-Foot HST 

Run 

28 

28 

28 

25 

25 

25 

Point 

1808 

1808 

1808 

1110 

1110 

1110 

M∞ 

0.50 

0.50 

0.50 

0.50 

0.50 

0.50 

Re/ft (10-6) 

2.5 

2.5 

2.5 

3.0 

3.0 

3.0 

α, deg 

20.02 

20.02 

20.02 

20.16 

20.16 

20.16 

x/c 

0.30 

0.60 

0.80 

0.30 

0.60 

0.80 

LEX 

Off 

Off 

Off 

Off 

Off 

Off 

Vertical Tail 

Twin δv=30o

Twin δv=30o

Twin δv=30o

Twin δv=30o

Twin δv=30o

Twin δv=30o

plot_dq1057d 

Thu Jul 15 21:53:41 2010 

-.16 

-.12 

-.08 

-.04 

0 

.04 

.08 

.12 

ΔCp 

ΔCp = Cp(initial point) - Cp(repeat point) ESP accuracy = 0.05% F.S. or ΔCp = +/-0.0038 at M = 0.50 

 

(b) 20 degrees angle of attack 

 Figure 47. Concluded.

 148
-4 0 4 8 12 16 20 24 28 32 36 40 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

LEX 

Solid 

Solid 

Re/ft (10-6) 

2.5 

3.0 

Vertical Tail 

Tail Off 

Tail Off 

 

Figure 48. Tunnel-to-tunnel comparison of the longitudinal aerodynamic characteristics at Mach = 0.50 

 with solid LEX; tail off.

 149
-4 0 4 8 12 16 20 24 28 32 36 40 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

LEX 

Porous 

Porous 

Re/ft (10-6) 

2.5 

3.0 

Vertical Tail 

CL Tail 

Tail Off 

 

Figure 49. Tunnel-to-tunnel comparison of the longitudinal aerodynamic characteristics at Mach = 0.50 

 with porous LEX; centerline tail on (T1057), tail off (T202).

 150
-4 0 4 8 12 16 20 24 28 32 36 40 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

LEX 

Off 

Off 

Re/ft (10-6) 

2.5 

3.0 

Vertical Tail 

CL Tail 

Tail Off 

 

Figure 50. Tunnel-to-tunnel comparison of the longitudinal aerodynamic characteristics at Mach = 0.50 

 with wing alone; centerline tail on (T1057), tail off (T202).

 151
-4 0 4 8 12 16 20 24 28 32 36 40 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

LEX 

Solid 

Solid 

Re/ft (10-6) 

2.5 

3.0 

Vertical Tail 

Twin δv=0o

Twin δv=0o

 

Figure 51. Tunnel-to-tunnel comparison of the longitudinal aerodynamic characteristics at Mach = 0.50 

 with solid LEX and twin uncanted tails.

 152
-4 0 4 8 12 16 20 24 28 32 36 40 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

LEX 

Solid 

Solid 

Re/ft (10-6) 

2.5 

3.0 

Vertical Tail 

Twin δv=30o

Twin δv=30o

 

Figure 52. Tunnel-to-tunnel comparison of the longitudinal aerodynamic characteristics at Mach = 0.50 

 with solid LEX and twin canted tails.

 153
-4 0 4 8 12 16 20 24 28 32 36 40 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

LEX 

Porous 

Porous 

Re/ft (10-6) 

2.5 

3.0 

Vertical Tail 

Twin δv=0o

Twin δv=0o

 

Figure 53. Tunnel-to-tunnel comparison of the longitudinal aerodynamic characteristics at Mach = 0.50 

 with porous LEX and twin uncanted tails.

 154
-4 0 4 8 12 16 20 24 28 32 36 40 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

LEX 

Off 

Off 

Re/ft (10-6) 

2.5 

3.0 

Vertical Tail 

Twin δv=30o

Twin δv=30o

 

Figure 54. Tunnel-to-tunnel comparison of the longitudinal aerodynamic characteristics at Mach = 0.50 

 with wing alone and twin canted tails.

 155
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.8 

-.4 

0 

.4 

.8 

1.2 

1.6 

2.0 

CN 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

23. 

8. 

Polar 

83. 

1. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Solid 

Solid 

Vertical Tail 

Tail Off 

Tail Off 

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Mon Jul 12 21:48:58 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.010 

-.005 

0 

.005 

.010 

ΔCN 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (a) normal force coefficient 

Figure 55. Tunnel-to-tunnel comparison of the delta longitudinal aerodynamic coefficients at Mach = 0.50 

 with solid LEX and tail off.

 156
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.004 

0 

.004 

.008 

.012 

.016 

.020 

.024 

CA 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

23. 

8. 

Polar 

83. 

1. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Solid 

Solid 

Vertical Tail 

Tail Off 

Tail Off 

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Mon Jul 12 21:55:21 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCA 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

(b) axial force coefficient 

 Figure 55. Continued.

 157
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

0 

.02 

.04 

.06 

.08 

.10 

.12 

.14 

Cm 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

23. 

8. 

Polar 

83. 

1. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Solid 

Solid 

Vertical Tail 

Tail Off 

Tail Off 

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Mon Jul 12 21:57:23 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.004 

-.002 

0 

.002 

.004 

ΔCm 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (c) pitching moment coefficient 

 Figure 55. Concluded.

 158
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.8 

-.4 

0 

.4 

.8 

1.2 

1.6 

2.0 

CN 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

20. 

15. 

Polar 

70. 

2. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Solid 

Solid 

Vertical Tail 

Twin δv=0o

Twin δv=0o

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Tue Jul 13 08:49:31 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.02 

-.01 

0 

.01 

.02 

ΔCN 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (a) normal force coefficient 

Figure 56. Tunnel-to-tunnel comparison of the delta longitudinal aerodynamic coefficients at Mach = 0.50 

 with solid LEX and twin uncanted tails.

 159
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.004 

0 

.004 

.008 

.012 

.016 

.020 

.024 

CA 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

20. 

15. 

Polar 

70. 

2. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Solid 

Solid 

Vertical Tail 

Twin δv=0o

Twin δv=0o

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Tue Jul 13 08:54:34 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCA 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

(b) axial force coefficient 

 Figure 56. Continued.

 160
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

0 

.04 

.08 

.12 

.16 

.20 

.24 

.28 

Cm 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

20. 

15. 

Polar 

70. 

2. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Solid 

Solid 

Vertical Tail 

Twin δv=0o

Twin δv=0o

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Tue Jul 13 10:20:26 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.004 

-.002 

0 

.002 

.004 

ΔCm 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (c) pitching moment coefficient 

 Figure 56. Concluded.

 161
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.8 

-.4 

0 

.4 

.8 

1.2 

1.6 

2.0 

CN 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

26. 

5. 

Polar 

94. 

1. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Solid 

Solid 

Vertical Tail 

Twin δv=30o

Twin δv=30o

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Tue Jul 13 08:10:55 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.02 

-.01 

0 

.01 

.02 

ΔCN 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (a) normal force coefficient 

Figure 57. Tunnel-to-tunnel comparison of the delta longitudinal aerodynamic coefficients at Mach = 0.50 

 with solid LEX and twin canted tails.

 162
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.004 

0 

.004 

.008 

.012 

.016 

.020 

.024 

CA 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

26. 

5. 

Polar 

94. 

1. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Solid 

Solid 

Vertical Tail 

Twin δv=30o

Twin δv=30o

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Tue Jul 13 08:18:08 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCA 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

(b) axial force coefficient 

 Figure 57. Continued.

 163
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

0 

.04 

.08 

.12 

.16 

.20 

.24 

.28 

Cm 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

26. 

5. 

Polar 

94. 

1. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Solid 

Solid 

Vertical Tail 

Twin δv=30o

Twin δv=30o

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Tue Jul 13 08:24:53 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.008 

-.004 

0 

.004 

.008 

ΔCm 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (c) pitching moment coefficient 

 Figure 57. Concluded.

 164
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.8 

-.4 

0 

.4 

.8 

1.2 

1.6 

2.0 

CN 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

17. 

14. 

Polar 

48. 

1. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Porous 

Porous 

Vertical Tail 

Twin δv=0o

Twin δv=0o

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Tue Jul 13 09:40:30 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.04 

-.02 

0 

.02 

.04 

ΔCN 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (a) normal force coefficient 

Figure 58. Tunnel-to-tunnel comparison of the delta longitudinal aerodynamic coefficients at Mach = 0.50 

 with porous LEX and twin uncanted tails.

 165
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

.012 

.016 

.020 

.024 

.028 

.032 

.036 

.040 

CA 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

17. 

14. 

Polar 

48. 

1. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Porous 

Porous 

Vertical Tail 

Twin δv=0o

Twin δv=0o

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Tue Jul 13 09:46:40 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCA 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

(b) axial force coefficient 

 Figure 58. Continued.

 166
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

0 

.02 

.04 

.06 

.08 

.10 

.12 

.14 

Cm 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

17. 

14. 

Polar 

48. 

1. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Porous 

Porous 

Vertical Tail 

Twin δv=0o

Twin δv=0o

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Tue Jul 13 09:52:56 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.004 

-.002 

0 

.002 

.004 

ΔCm 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (c) pitching moment coefficient 

 Figure 58. Concluded.

 167
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.8 

-.4 

0 

.4 

.8 

1.2 

1.6 

2.0 

CN 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

28. 

25. 

Polar 

107. 

1. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Off 

Off 

Vertical Tail 

Twin δv=30o

Twin δv=30o

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Tue Jul 13 10:02:37 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.02 

-.01 

0 

.01 

.02 

ΔCN 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (a) normal force coefficient 

Figure 59. Tunnel-to-tunnel comparison of the delta longitudinal aerodynamic coefficients at Mach = 0.50 

 with wing alone and twin canted tails.

 168
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.008 

-.004 

0 

.004 

.008 

.012 

.016 

.020 

CA 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

28. 

25. 

Polar 

107. 

1. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Off 

Off 

Vertical Tail 

Twin δv=30o

Twin δv=30o

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Tue Jul 13 10:09:37 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.002 

-.001 

0 

.001 

.002 

ΔCA 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

(b) axial force coefficient 

 Figure 59. Continued.

 169
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

.10 

Cm 

Test No. 

1057. 

202. 

Facility 

8-Foot TPT 

7x10-Foot HST 

Run 

28. 

25. 

Polar 

107. 

1. 

M∞ 

0.50 

0.50 

Re/ft (10-6) 

2.5 

3.0 

LEX 

Off 

Off 

Vertical Tail 

Twin δv=30o

Twin δv=30o

 Tests 1057 and 202 Porous LEX Model NASA LaRC 8-Foot TPT and 7x10-Foot HST 

Tue Jul 13 10:44:33 2010 

Δ’s are obtained by interpolating in each polar to the nominal values of the independent 

variable, then averaging and subtracting the averages from the interpolated data. 

-.004 

-.002 

0 

.002 

.004 

ΔCm 

Dashed lines computed from NASA LaRC 755 balance accuracy (95-percent confidence) 

 (c) pitching moment coefficient 

 Figure 59. Concluded.

 170
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (a) 12 degrees angle of attack Figure 60. Effect of the Mach number on the wing upper surface static pressure distributions with wing alone and 
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 16 degrees angle of attack  Figure 60. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (c) 20 degrees angle of attack  Figure 60. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off Off Off Off 

α, deg 

8.00 

12.01 16.20 20.03 24.01 28.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 Figure 61. Effect of the angle of attack on the wing upper surface static pressure distributions at Mach = 0.50 with wing a
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off Off Off Off 

α, deg 

8.00 

12.02 16.04 19.99 24.01 28.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 Figure 62. Effect of the angle of attack on the wing upper surface static pressure distributions at Mach = 0.85 with wing a
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off Off Off 

α, deg 

7.98 

11.99 16.01 19.96 22.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 Figure 63. Effect of the angle of attack on the wing upper surface static pressure distributions at Mach = 1.20 with wing a
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off Off Off 

β, deg 

-7.99 -4.00 -0.01 

4.02 7.81 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (a) 12 degrees angle of attack Figure 64. Effect of the sideslip angle on the wing upper surface static pressure distributions at Mach = 0.50 with wing al
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off Off Off 

β, deg 

-8.01 -3.99 -0.01 

4.05 7.66 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 16 degrees angle of attack  Figure 64. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off Off Off 

β, deg 

-8.01 -4.06 -0.02 

4.00 7.45 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (c) 20 degrees angle of attack  Figure 64. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off Off Off 

β, deg 

-7.99 -4.02 -0.01 

4.00 7.26 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (d) 24 degrees angle of attack  Figure 64. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off Off Off 

β, deg 

-8.06 -3.99 -0.02 

3.98 7.71 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (a) 12 degrees angle of attack Figure 65. Effect of the sideslip angle on the wing upper surface static pressure distributions at Mach = 0.85 with wing al
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off Off Off 

β, deg 

-8.04 -4.02 

0.01 4.01 7.71 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 16 degrees angle of attack  Figure 65. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off Off Off 

β, deg 

-8.01 -4.06 

0.01 4.08 7.56 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (c) 20 degrees angle of attack  Figure 65. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off Off Off 

β, deg 

-8.15 -4.04 

0.02 4.11 7.50 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (d) 24 degrees angle of attack  Figure 65. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off Off Off 

β, deg 

-7.95 -4.01 -0.06 

4.02 7.64 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (a) 12 degrees angle of attack Figure 66. Effect of the sideslip angle on the wing upper surface static pressure distributions at Mach = 1.20 with wing al
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Off Off Off Off Off 

β, deg 

-8.00 -4.04 

0.08 4.14 7.56 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 16 degrees angle of attack  Figure 66. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 12 degrees angle of attack Figure 67. Effect of the Mach number on the wing upper surface static pressure distributions with solid LEX and 
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 16 degrees angle of attack  Figure 67. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (c) 20 degrees angle of attack  Figure 67. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (d) 24 degrees angle of attack  Figure 67. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid Solid 

α, deg 

8.01 

12.01 16.04 20.03 24.06 28.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 Figure 68. Effect of the angle of attack on the wing upper surface static pressure distributions at Mach = 0.50 with solid L
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid Solid 

α, deg 

8.08 

12.02 16.12 20.13 24.09 27.98 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 Figure 69. Effect of the angle of attack on the wing upper surface static pressure distributions at Mach = 0.85 with solid L
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid 

α, deg 

8.01 

12.00 16.03 20.04 24.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 Figure 70. Effect of the angle of attack on the wing upper surface static pressure distributions at Mach = 1.20 with solid L
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid 

β, deg 

-8.00 -3.95 

0.02 4.06 7.78 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(a) 12 degrees angle of attack 

Figure 71. Effect of the sideslip angle on the wing upper surface static pressure distributions at Mach = 0.50 with solid L
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid 

β, deg 

-8.10 -4.00 

0.02 3.98 7.59 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(b) 16 degrees angle of attack 

Figure 71. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid 

β, deg 

-7.99 -4.03 

0.00 4.10 7.57 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(c) 20 degrees angle of attack 

Figure 71. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid 

β, deg 

-7.95 -4.00 

0.00 4.09 7.24 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(d) 24 degrees angle of attack 

Figure 71. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid 

β, deg 

-8.03 -4.03 

0.00 4.05 7.82 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

a) 12 degrees angle of attack 

Figure 72. Effect of the sideslip angle on the wing upper surface static pressure distributions at Mach = 0.85 with solid L
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid 

β, deg 

-8.03 -4.02 -0.01 

4.04 7.70 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(b) 16 degrees angle of attack 

Figure 72. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid 

β, deg 

-8.05 -3.99 

0.00 4.01 7.58 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(c) 20 degrees angle of attack 

Figure 72. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid 

β, deg 

-7.52 -4.08 

0.01 4.04 7.24 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(d) 24 degrees angle of attack 

Figure 72. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid 

β, deg 

-8.01 -4.00 

0.00 4.03 7.67 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(a) 12 degrees angle of attack 

Figure 73. Effect of the sideslip angle on the wing upper surface static pressure distributions at Mach = 1.20 with solid L
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid 

β, deg 

-7.99 -3.99 

0.00 4.01 7.65 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(b) 16 degrees angle of attack 

Figure 73. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid Solid Solid Solid Solid 

β, deg 

-8.00 -4.00 

0.00 4.02 7.54 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(c) 20 degrees angle of attack 

Figure 73. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 12 degrees angle of attack Figure 74. Effect of the Mach number on the wing upper surface static pressure distributions with porous LEX and 
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 16 degrees angle of attack  Figure 74. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (c) 20 degrees angle of attack  Figure 74. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (d) 24 degrees angle of attack  Figure 74. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous Porous 

α, deg 

8.00 

12.01 16.08 20.05 24.04 28.08 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 Figure 75. Effect of the angle of attack on the wing upper surface static pressure distributions at Mach = 0.50 with porous 
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous Porous 

α, deg 

8.03 

12.05 16.00 20.03 24.02 28.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 Figure 76. Effect of the angle of attack on the wing upper surface static pressure distributions at Mach = 0.85 with porous 
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous 

α, deg 

7.99 

12.01 16.00 20.01 24.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 Figure 77. Effect of the angle of attack on the wing upper surface static pressure distributions at Mach = 1.20 with porous 
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous 

β, deg 

-8.00 -4.00 -0.01 

4.02 7.75 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

  (a) 12 degrees angle of attack Figure 78. Effect of the sideslip angle on the wing upper surface static pressure distributions at Mach = 0.50 with porous L
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous 

β, deg 

-8.00 -4.01 -0.01 

4.00 7.63 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

  (b) 16 degrees angle of attack  Figure 78. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous 

β, deg 

-8.00 -3.99 

0.00 4.03 7.52 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

  (c) 20 degrees angle of attack  Figure 78. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous 

β, deg 

-8.00 -4.00 

0.00 4.04 7.22 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

  (d) 24 degrees angle of attack  Figure 78. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous 

β, deg 

-8.02 -4.02 

0.00 4.01 7.75 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

  (a) 12 degrees angle of attack Figure 79. Effect of the sideslip angle on the wing upper surface static pressure distributions at Mach = 0.85 with porous L
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous 

β, deg 

-8.00 -4.00 -0.01 

4.02 7.62 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

  (b) 16 degrees angle of attack  Figure 79. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous 

β, deg 

-8.01 -4.03 

0.00 4.01 7.42 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

  (c) 20 degrees angle of attack  Figure 79. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous 

β, deg 

-8.05 -4.01 

0.01 4.01 7.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

  (d) 24 degrees angle of attack  Figure 79. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous 

β, deg 

-8.03 -4.02 

0.01 4.01 7.73 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

  (a) 12 degrees angle of attack Figure 80. Effect of the sideslip angle on the wing upper surface static pressure distributions at Mach = 1.20 with porous L
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous 

β, deg 

-8.10 -4.00 

0.01 4.00 7.66 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

  (b) 16 degrees angle of attack  Figure 80. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Porous Porous Porous Porous Porous 

β, deg 

-7.98 -4.01 

0.01 4.01 7.42 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

  (c) 20 degrees angle of attack  Figure 80. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

 (a) 12 degrees angle of attack Figure 81. Effect of the Mach number on the wing upper surface static pressure distributions with canard and c
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

 (b) 16 degrees angle of attack  Figure 81. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

 (c) 20 degrees angle of attack  Figure 81. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On 

M∞ 

0.50 0.85 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

 (d) 24 degrees angle of attack  Figure 81. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On On 

α, deg 

8.05 

12.03 16.05 20.05 24.04 28.05 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

 Figure 82. Effect of the angle of attack on the wing upper surface static pressure distributions at Mach = 0.50 with cana
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On On 

α, deg 

8.02 

12.03 16.05 20.06 24.04 28.06 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

 Figure 83. Effect of the angle of attack on the wing upper surface static pressure distributions at Mach = 0.85 with cana
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On 

α, deg 

8.07 

12.02 16.02 20.03 24.05 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

 Figure 84. Effect of the angle of attack on the wing upper surface static pressure distributions at Mach = 1.20 with cana
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On 

β, deg 

-8.07 -4.08 -0.01 

4.00 7.75 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

  (a) 12 degrees angle of attack Figure 85. Effect of the sideslip angle on the wing upper surface static pressure distributions at Mach = 0.50 with canar
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On 

β, deg 

-8.01 -4.08 

0.02 4.03 7.54 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

  (b) 16 degrees angle of attack  Figure 85. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On 

β, deg 

-8.02 -4.01 

0.02 4.03 7.57 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

  (c) 20 degrees angle of attack  Figure 85. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On 

β, deg 

-7.93 -4.00 -0.01 

4.00 7.18 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

  (d) 24 degrees angle of attack  Figure 85. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On 

β, deg 

-8.12 -4.05 -0.02 

4.01 7.69 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

  (a) 12 degrees angle of attack Figure 86. Effect of the sideslip angle on the wing upper surface static pressure distributions at Mach = 0.85 with canar
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On 

β, deg 

-8.00 -4.04 

0.01 4.06 7.61 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

  (b) 16 degrees angle of attack  Figure 86. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On 

β, deg 

-7.99 -4.00 -0.01 

4.01 7.52 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

  (c) 20 degrees angle of attack  Figure 86. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On 

β, deg 

-7.99 -3.99 

0.00 4.00 7.29 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

  (d) 24 degrees angle of attack  Figure 86. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On 

β, deg 

-8.04 -4.07 

0.00 4.02 7.82 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

  (a) 12 degrees angle of attack Figure 87. Effect of the sideslip angle on the wing upper surface static pressure distributions at Mach = 1.20 with canar
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On 

β, deg 

-8.00 -4.06 -0.01 

4.03 7.82 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

  (b) 16 degrees angle of attack  Figure 87. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Canard 

On On On On On 

β, deg 

-8.01 -3.99 -0.01 

4.04 7.57 

x/c=0.80 

x/c=0.60 

x/c=0.30 

Canard →

  (c) 20 degrees angle of attack  Figure 87. Concluded.
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

LEX 

Off 

Off 

Off 

M∞ 

0.50 

0.85 

1.20 

 

Figure 88. Mach number effect on the longitudinal aerodynamic characteristics 

 with wing alone and centerline tail.

 241
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Off 

Off 

Off 

Off 

Vertical Tail 

Centerline 

Centerline 

Centerline 

Centerline 

α, deg 

12.01 

16.01 

20.02 

24.02 

 

Figure 89. Lateral-directional aerodynamic characteristics in sideslip sweeps at Mach = 0.50 

 with wing alone and centerline tail.

 242
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Off 

Off 

Off 

Off 

Vertical Tail 

Centerline 

Centerline 

Centerline 

Centerline 

α, deg 

12.00 

16.01 

20.00 

24.00 

 

Figure 90. Lateral-directional aerodynamic characteristics in sideslip sweeps at Mach = 0.85 

 with wing alone and centerline tail.

 243
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Off 

Off 

Vertical Tail 

Centerline 

Centerline 

α, deg 

12.06 

15.98 

 

Figure 91. Lateral-directional aerodynamic characteristics in sideslip sweeps at Mach = 1.20 

 with wing alone and centerline tail.

 244
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

LEX 

Solid 

Solid 

Solid 

M∞ 

0.50 

0.85 

1.20 

 

Figure 92. Mach number effect on the longitudinal aerodynamic characteristics 

 with solid LEX and centerline tail.

 245
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline 

Centerline 

Centerline 

Centerline 

Centerline 

α, deg 

8.06 

12.06 

16.13 

20.13 

24.15 

 

Figure 93. Lateral-directional aerodynamic characteristics in sideslip sweeps at Mach = 0.50 

 with solid LEX and centerline tail.

 246
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Solid 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline 

Centerline 

Centerline 

Centerline 

Centerline 

α, deg 

8.05 

12.05 

16.07 

20.07 

24.07 

 

Figure 94. Lateral-directional aerodynamic characteristics in sideslip sweeps at Mach = 0.85 

 with solid LEX and centerline tail.

 247
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Solid 

Solid 

Solid 

Solid 

Vertical Tail 

Centerline 

Centerline 

Centerline 

Centerline 

α, deg 

8.00 

12.01 

16.01 

20.01 

 

Figure 95. Lateral-directional aerodynamic characteristics in sideslip sweeps at Mach = 1.20 

 with solid LEX and centerline tail.

 248
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

LEX 

Porous 

Porous 

Porous 

M∞ 

0.50 

0.85 

1.20 

 

Figure 96. Mach number effect on the longitudinal aerodynamic characteristics 

 with porous LEX and centerline tail.

 249
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline 

Centerline 

Centerline 

Centerline 

Centerline 

α, deg 

8.03 

12.03 

16.04 

20.04 

24.04 

 

Figure 97. Lateral-directional aerodynamic characteristics in sideslip sweeps at Mach = 0.50 

 with porous LEX and centerline tail.

 250
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Porous 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline 

Centerline 

Centerline 

Centerline 

Centerline 

α, deg 

8.02 

12.02 

16.04 

20.04 

24.06 

 

Figure 98. Lateral-directional aerodynamic characteristics in sideslip sweeps at Mach = 0.85 

 with porous LEX and centerline tail.

 251
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Porous 

Porous 

Porous 

Porous 

Vertical Tail 

Centerline 

Centerline 

Centerline 

Centerline 

α, deg 

8.01 

12.00 

16.00 

20.01 

 

Figure 99. Lateral-directional aerodynamic characteristics in sideslip sweeps at Mach = 1.20 

 with porous LEX and centerline tail.

 252
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

Canard 

On 

On 

On 

M∞ 

0.50 

0.85 

1.20 

 

Figure 100. Mach number effect on the longitudinal aerodynamic characteristics 

 with canard and centerline tail.

 253
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Canard 

On 

On 

On 

On 

On 

Vertical Tail 

Centerline 

Centerline 

Centerline 

Centerline 

Centerline 

α, deg 

8.02 

12.02 

16.04 

20.04 

24.04 

 

Figure 101. Lateral-directional aerodynamic characteristics in sideslip sweeps at Mach = 0.50 

 with canard and centerline tail.

 254
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Canard 

On 

On 

On 

On 

On 

Vertical Tail 

Centerline 

Centerline 

Centerline 

Centerline 

Centerline 

α, deg 

8.02 

12.03 

16.04 

20.04 

24.05 

 

Figure 102. Lateral-directional aerodynamic characteristics in sideslip sweeps at Mach = 0.85 

 with canard and centerline tail.

 255
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Canard 

On 

On 

On 

On 

Vertical Tail 

Centerline 

Centerline 

Centerline 

Centerline 

α, deg 

8.05 

12.03 

16.04 

20.05 

 

Figure 103. Lateral-directional aerodynamic characteristics in sideslip sweeps at Mach = 1.20 

 with canard and centerline tail.

 256
 Figure 104. Laser vapor screen flow visualization images at Mach = 0.85 and x/c = 0.80 on the solid LEX, porous LEX

Solid LEX Porous LEX Wing AloSolid LEX Porous LEX Wing Alo

(a) 16 degrees angle of attack (b) 20 degrees angle of attack 
Solid LEX Porous LEX Wing Alo

(c) 24 degrees angle of attack (d) 28 degrees angle of attack 

Solid LEX Porous LEX Wing Alo

Figure 104. Concluded. 
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.50 0.50 0.50 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 8 degrees angle of attack Figure 105. Effect of LEX porosity on the wing upper surface static pressure distributions at Mach = 0.50 with c (Wing-alone data are shown for reference.)
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.50 0.50 0.50 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 10 degrees angle of attack  Figure 105. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.50 0.50 0.50 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (c) 12 degrees angle of attack  Figure 105. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.50 0.50 0.50 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (d) 14 degrees angle of attack  Figure 105. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.50 0.50 0.50 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (e) 16 degrees angle of attack  Figure 105. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.50 0.50 0.50 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (f) 18 degrees angle of attack  Figure 105. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.50 0.50 0.50 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (g) 20 degrees angle of attack  Figure 105. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.50 0.50 0.50 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (h) 22 degrees angle of attack  Figure 105. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.50 0.50 0.50 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (i) 24 degrees angle of attack  Figure 105. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.50 0.50 0.50 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (j) 26 degrees angle of attack  Figure 105. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.50 0.50 0.50 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (f) 28 degrees angle of attack  Figure 105. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

12.17 12.04 12.00 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 12 degrees angle of attack, -8 degrees angle of sideslip Figure 106. Effect of LEX porosity on the wing upper surface static pressure distributions in sideslip sweeps at Mach = 0 (Wing-alone data are shown for reference.)
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

12.06 12.01 12.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 12 degrees angle of attack, -4 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

12.04 12.05 12.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (c) 12 degrees angle of attack, +4 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

12.14 12.07 12.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (d) 12 degrees angle of attack, +8 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

16.30 16.00 16.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (e) 16 degrees angle of attack, -8 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

16.12 16.04 16.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (f) 16 degrees angle of attack, -4 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

16.09 16.03 16.00 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (g) 16 degrees angle of attack, +4 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

16.22 16.10 16.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (h) 16 degrees angle of attack, +8 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

20.33 20.04 20.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (i) 20 degrees angle of attack, -8 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

20.12 20.03 20.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (j) 20 degrees angle of attack, -4 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

20.09 20.06 20.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (k) 20 degrees angle of attack, +4 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

20.25 20.09 20.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (l) 20 degrees angle of attack, +8 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

24.37 24.02 24.00 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (m) 24 degrees angle of attack, -8 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

24.13 24.02 24.07 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (n) 24 degrees angle of attack, -4 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

24.10 24.05 24.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (o) 24 degrees angle of attack, +4 degrees angle of sideslip  Figure 106. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

24.27 24.08 24.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (p) 24 degrees angle of attack, +8 degrees angle of sideslip  Figure 106. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.85 0.85 0.85 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 8 degrees angle of attack Figure 107. Effect of LEX porosity on the wing upper surface static pressure distributions at Mach = 0.85 with c (Wing-alone data are shown for reference.)
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.85 0.85 0.85 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 10 degrees angle of attack  Figure 107. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.85 0.85 0.85 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (c) 12 degrees angle of attack  Figure 107. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.85 0.85 0.85 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (d) 14 degrees angle of attack  Figure 107. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.85 0.85 0.85 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (e) 16 degrees angle of attack  Figure 107. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.85 0.85 0.85 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (f) 18 degrees angle of attack  Figure 107. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.85 0.85 0.85 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (g) 20 degrees angle of attack  Figure 107. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.85 0.85 0.85 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (h) 22 degrees angle of attack  Figure 107. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.85 0.85 0.85 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (i) 24 degrees angle of attack  Figure 107. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.85 0.85 0.85 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (j) 26 degrees angle of attack  Figure 107. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

0.85 0.85 0.85 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (k) 28 degrees angle of attack  Figure 107. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

12.01 12.00 11.99 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 12 degrees angle of attack, -8 degrees angle of sideslip Figure 108. Effect of LEX porosity on the wing upper surface static pressure distributions in sideslip sweeps at Mach = 0 (Wing-alone data are shown for reference.)
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

12.07 12.02 12.00 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 12 degrees angle of attack, -4 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

12.05 12.02 12.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (c) 12 degrees angle of attack, +4 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

12.05 12.07 12.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (d) 12 degrees angle of attack, +8 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

16.08 16.02 15.98 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (e) 16 degrees angle of attack, -8 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

16.04 16.05 15.99 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (f) 16 degrees angle of attack, -4 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

16.09 16.04 16.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (g) 16 degrees angle of attack, +4 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

16.05 16.07 16.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (h) 16 degrees angle of attack, +8 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

20.05 20.04 19.97 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (i) 20 degrees angle of attack, -8 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

20.11 20.05 20.00 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (j) 20 degrees angle of attack, -4 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

20.12 20.04 20.00 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (k) 20 degrees angle of attack, +4 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

20.08 20.14 20.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (l) 20 degrees angle of attack, +8 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

24.08 24.01 24.00 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (m) 24 degrees angle of attack, -8 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

24.11 24.10 24.00 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (n) 24 degrees angle of attack, -4 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

24.10 24.04 23.99 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (o) 24 degrees angle of attack, +4 degrees angle of sideslip  Figure 108. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

24.08 24.03 23.99 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (p) 24 degrees angle of attack, +8 degrees angle of sideslip  Figure 108. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

1.20 1.20 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 8 degrees angle of attack Figure 109. Effect of LEX porosity on the wing upper surface static pressure distributions at Mach = 1.20 with c (Wing-alone data are shown for reference.)
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

1.20 1.20 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 10 degrees angle of attack  Figure 109. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

1.20 1.20 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (c) 12 degrees angle of attack  Figure 109. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

1.20 1.20 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (d) 14 degrees angle of attack  Figure 109. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

1.20 1.20 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (e) 16 degrees angle of attack  Figure 109. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

1.20 1.20 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (f) 18 degrees angle of attack  Figure 109. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

1.20 1.20 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (g) 20 degrees angle of attack  Figure 109. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

M∞ 

1.20 1.20 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (h) 22 degrees angle of attack  Figure 109. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

M∞ 

1.20 1.20 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (i) 24 degrees angle of attack  Figure 109. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

12.00 11.99 12.06 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 12 degrees angle of attack, -8 degrees angle of sideslip Figure 110. Effect of LEX porosity on the wing upper surface static pressure distributions in sideslip sweeps at Mach = 1 (Wing-alone data are shown for reference.)
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

12.01 11.99 12.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 12 degrees angle of attack, -4 degrees angle of sideslip  Figure 110. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

12.00 12.02 12.09 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (c) 12 degrees angle of attack, +4 degrees angle of sideslip  Figure 110. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

12.02 12.01 12.08 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (d) 12 degrees angle of attack, +8 degrees angle of sideslip  Figure 110. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

16.01 15.99 15.97 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (e) 16 degrees angle of attack, -8 degrees angle of sideslip  Figure 110. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

16.01 16.00 15.98 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (f) 16 degrees angle of attack, -4 degrees angle of sideslip  Figure 110. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

16.01 16.00 15.99 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (g) 16 degrees angle of attack, +4 degrees angle of sideslip  Figure 110. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX 

Solid 

Porous 

Off 

α, deg 

16.02 16.00 15.98 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (h) 16 degrees angle of attack, +8 degrees angle of sideslip  Figure 110. Concluded.
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

LEX 

Solid 

Porous 

Off 

 (a) lift, drag, and pitching moment coefficients 

Figure 111. Effect of LEX porosity on the six-component aerodynamic characteristics at Mach = 0.50 

 with centerline tail. (Wing-alone data are shown for reference.)

 330
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

CY 

-.016 

-.012 

-.008 

-.004 

0 

.004 

.008 

.012 

.016 

Cn 

-.008 

-.006 

-.004 

-.002 

0 

.002 

.004 

.006 

.008 

Cl 

LEX 

Solid 

Porous 

Off 

 (b) rolling moment, yawing moment, and side force coefficients 

 Figure 111. Concluded.

 331
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

LEX 

Solid 

Porous 

Off 

 (a) lift, drag, and pitching moment coefficients 

Figure 112. Effect of LEX porosity on the six-component aerodynamic characteristics at Mach = 0.85 

 with centerline tail. (Wing-alone data are shown for reference.)

 332
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

CY 

-.016 

-.012 

-.008 

-.004 

0 

.004 

.008 

.012 

.016 

Cn 

-.004 

0 

.004 

.008 

.012 

.016 

.020 

.024 

.028 

Cl 

LEX 

Solid 

Porous 

Off 

 (b) rolling moment, yawing moment, and side force coefficients 

 Figure 112. Concluded.

 333
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

LEX 

Solid 

Porous 

Off 

 (a) lift, drag, and pitching moment coefficients 

Figure 113. Effect of LEX porosity on the six-component aerodynamic characteristics at Mach = 1.20 

 with centerline tail. (Wing-alone data are shown for reference.)

 334
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

CY 

-.016 

-.012 

-.008 

-.004 

0 

.004 

.008 

.012 

.016 

Cn 

-.008 

-.006 

-.004 

-.002 

0 

.002 

.004 

.006 

.008 

Cl 

LEX 

Solid 

Porous 

Off 

 (b) rolling moment, yawing moment, and side force coefficients 

 Figure 113. Concluded.

 335
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

0 

.01 

.02 

.03 

.04 

.05 

.06 

.07 

.08 

Cm 

.08 

.09 

.10 

.11 

.12 

.13 

.14 

.15 

.16 

CD 

.44 

.46 

.48 

.50 

.52 

.54 

.56 

.58 

.60 

CL 

LEX 

Solid 

Porous 

Off 

 (a) lift, drag, and pitching moment coefficients at 12 degrees angle of attack 

Figure 114. Effect of LEX porosity on the six-component aerodynamic characteristics in sideslip sweeps 

 at Mach = 0.50 with centerline tail. (Wing-alone data are shown for reference.)

 336
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Solid 

Porous 

Off 

(b) rolling moment, yawing moment, and side force coefficients at 12 degrees angle of attack 

 Figure 114. Continued.

 337
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

0 

.01 

.02 

.03 

.04 

.05 

.06 

.07 

.08 

Cm 

.18 

.19 

.20 

.21 

.22 

.23 

.24 

.25 

.26 

CD 

.64 

.66 

.68 

.70 

.72 

.74 

.76 

.78 

.80 

CL 

LEX 

Solid 

Porous 

Off 

(c) lift, drag, and pitching moment coefficients at 16 degrees angle of attack 

 Figure 114. Continued.

 338
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Solid 

Porous 

Off 

(d) rolling moment, yawing moment, and side force coefficients at 16 degrees angle of attack 

 Figure 114. Continued.

 339
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

0 

.01 

.02 

.03 

.04 

.05 

.06 

.07 

.08 

Cm 

.28 

.30 

.32 

.34 

.36 

.38 

.40 

.42 

.44 

CD 

.76 

.80 

.84 

.88 

.92 

.96 

1.00 

1.04 

1.08 

CL 

LEX 

Solid 

Porous 

Off 

(e) lift, drag, and pitching moment coefficients at 20 degrees angle of attack 

 Figure 114. Continued.

 340
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Solid 

Porous 

Off 

(f) rolling moment, yawing moment, and side force coefficients at 20 degrees angle of attack 

 Figure 114. Continued.

 341
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

0 

.02 

.04 

.06 

.08 

.10 

.12 

.14 

.16 

Cm 

.42 

.44 

.46 

.48 

.50 

.52 

.54 

.56 

.58 

CD 

.92 

.96 

1.00 

1.04 

1.08 

1.12 

1.16 

1.20 

1.24 

CL 

LEX 

Solid 

Porous 

Off 

(g) lift, drag, and pitching moment coefficients at 24 degrees angle of attack 

 Figure 114. Continued.

 342
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Solid 

Porous 

Off 

(h) rolling moment, yawing moment, and side force coefficients at 24 degrees angle of attack 

 Figure 114. Concluded.

 343
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cm 

.10 

.11 

.12 

.13 

.14 

.15 

.16 

.17 

.18 

CD 

.50 

.52 

.54 

.56 

.58 

.60 

.62 

.64 

.66 

CL 

LEX 

Solid 

Porous 

Off 

 (a) lift, drag, and pitching moment coefficients at 12 degrees angle of attack 

Figure 115. Effect of LEX porosity on the six-component aerodynamic characteristics in sideslip sweeps 

 at Mach = 0.85 with centerline tail. (Wing-alone data are shown for reference.)

 344
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Solid 

Porous 

Off 

(b) rolling moment, yawing moment, and side force coefficients at 12 degrees angle of attack 

 Figure 115. Continued.

 345
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cm 

.20 

.21 

.22 

.23 

.24 

.25 

.26 

.27 

.28 

CD 

© 

.70 

.72 

.74 

.76 

.78 

.80 

.82 

.84 

.86 

CL 

LEX 

Solid 

Porous 

Off 

(c) lift, drag, and pitching moment coefficients at 16 degrees angle of attack 

 Figure 115. Continued.

 346
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Solid 

Porous 

Off 

(d) rolling moment, yawing moment, and side force coefficients at 16 degrees angle of attack 

 Figure 115. Continued.

 347
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cm 

.32 

.33 

.34 

.35 

.36 

.37 

.38 

.39 

.40 

CD 

.84 

.86 

.88 

.90 

.92 

.94 

.96 

.98 

1.00 

CL 

LEX 

Solid 

Porous 

Off 

(e) lift, drag, and pitching moment coefficients at 20 degrees angle of attack 

 Figure 115. Continued.

 348
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Solid 

Porous 

Off 

(f) rolling moment, yawing moment, and side force coefficients at 20 degrees angle of attack 

 Figure 115. Continued.

 349
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cm 

.40 

.42 

.44 

.46 

.48 

.50 

.52 

.54 

.56 

CD 

.84 

.88 

.92 

.96 

1.00 

1.04 

1.08 

1.12 

1.16 

CL 

LEX 

Solid 

Porous 

Off 

(g) lift, drag, and pitching moment coefficients at 24 degrees angle of attack 

 Figure 115. Continued.

 350
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Solid 

Porous 

Off 

(h) rolling moment, yawing moment, and side force coefficients at 24 degrees angle of attack 

 Figure 115. Concluded.

 351
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.07 

-.06 

-.05 

-.04 

-.03 

-.02 

-.01 

0 

.01 

Cm 

.10 

.11 

.12 

.13 

.14 

.15 

.16 

.17 

.18 

CD 

.50 

.52 

.54 

.56 

.58 

.60 

.62 

.64 

.66 

CL 

LEX 

Solid 

Porous 

Off 

 (a) lift, drag, and pitching moment coefficients at 12 degrees angle of attack 

Figure 116. Effect of LEX porosity on the six-component aerodynamic characteristics in sideslip sweeps 

 at Mach = 1.20 with centerline tail. (Wing-alone data are shown for reference.)

 352
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Solid 

Porous 

Off 

(b) rolling moment, yawing moment, and side force coefficients at 12 degrees angle of attack 

 Figure 116. Continued.

 353
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.07 

-.06 

-.05 

-.04 

-.03 

-.02 

-.01 

0 

.01 

Cm 

.20 

.21 

.22 

.23 

.24 

.25 

.26 

.27 

.28 

CD 

.64 

.66 

.68 

.70 

.72 

.74 

.76 

.78 

.80 

CL 

LEX 

Solid 

Porous 

Off 

(c) lift, drag, and pitching moment coefficients at 16 degrees angle of attack 

 Figure 116. Continued.

 354
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

LEX 

Solid 

Porous 

Off 

(d) rolling moment, yawing moment, and side force coefficients at 16 degrees angle of attack 

 Figure 116. Concluded.

 355
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) Level 3 (5.4%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

(a) 8 degrees angle of attack 

Figure 117. Comparison of the wing upper surface static pressure distributions at Mach = 0.50 with LEX porosity levels 1, 2
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) Level 3 (5.4%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 12 degrees angle of attack  Figure 117. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) Level 3 (5.4%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (c) 16 degrees angle of attack  Figure 117. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) Level 3 (5.4%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (d) 20 degrees angle of attack  Figure 117. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) Level 3 (5.4%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (e) 24 degrees angle of attack  Figure 117. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) Level 3 (5.4%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (f) 28 degrees angle of attack  Figure 117. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) Level 3 (5.4%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

(a) 8 degrees angle of attack 

Figure 118. Comparison of the wing upper surface static pressure distributions at Mach = 0.85 with LEX porosity levels 1, 2
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) Level 3 (5.4%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 12 degrees angle of attack  Figure 118. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) Level 3 (5.4%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (c) 16 degrees angle of attack  Figure 118. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) Level 3 (5.4%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (d) 20 degrees angle of attack  Figure 118. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) Level 3 (5.4%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (e) 24 degrees angle of attack  Figure 118. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) Level 3 (5.4%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (f) 28 degrees angle of attack  Figure 118. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) Level 5 (8.7%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

(a) 8 degrees angle of attack 

Figure 119. Comparison of the wing upper surface static pressure distributions at Mach = 0.50 with LEX porosity levels 1, 4
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) Level 5 (8.7%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 12 degrees angle of attack  Figure 119. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) Level 5 (8.7%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (c) 16 degrees angle of attack  Figure 119. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) Level 5 (8.7%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (d) 18 degrees angle of attack  Figure 119. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) Level 5 (8.7%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (e) 24 degrees angle of attack  Figure 119. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) Level 5 (8.7%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (f) 28 degrees angle of attack  Figure 119. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) Level 5 (8.7%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

(a) 8 degrees angle of attack 

Figure 120. Comparison of the wing upper surface static pressure distributions at Mach = 0.85 with LEX porosity levels 1, 4
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) Level 5 (8.7%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 12 degrees angle of attack  Figure 120. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) Level 5 (8.7%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (c) 18 degrees angle of attack  Figure 120. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) Level 5 (8.7%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (d) 20 degrees angle of attack  Figure 120. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) Level 5 (8.7%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (e) 24 degrees angle of attack  Figure 120. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) Level 5 (8.7%) 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (f) 28 degrees angle of attack  Figure 120. Concluded.
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) 

Level 3 (5.4%) 

Figure 121. Comparison of the longitudinal aerodynamic characteristics with porosity levels 1, 2, and 3 

 at Mach = 0.50; centerline tail on.

 380
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

LEX Porosity 

Level 1 (14.8%) 

Level 2 (9.4%) 

Level 3 (5.4%) 

Figure 122. Comparison of the longitudinal aerodynamic characteristics with porosity levels 1, 2, and 3 

 at Mach = 0.85; centerline tail on.

 381
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) 

Level 5 (8.7%) 

Figure 123. Comparison of the longitudinal aerodynamic characteristics with porosity levels 1, 4, and 5 

 at Mach = 0.50; centerline tail on.

 382
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

LEX Porosity 

Level 1 (14.8%) 

Level 4 (6.1%) 

Level 5 (8.7%) 

Figure 124. Comparison of the longitudinal aerodynamic characteristics with porosity levels 1, 4, and 5 

 at Mach = 0.85; centerline tail on.

 383
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

8.01 8.03 7.99 8.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 8 degrees angle of attack Figure 125. Comparison of the wing upper surface static pressure distributions at Mach = 0.50 with centerline tail, tw twin canted tails, and tail off; solid LEX.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

12.01 12.08 12.01 12.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 12 degrees angle of attack  Figure 125. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

16.04 16.04 16.03 16.07 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (c) 16 degrees angle of attack  Figure 125. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

20.03 20.03 20.06 20.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (d) 20 degrees angle of attack  Figure 125. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

24.06 24.06 24.09 24.09 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (e) 24 degrees angle of attack  Figure 125. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

α, deg 

28.02 28.04 28.09 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (f) 28 degrees angle of attack  Figure 125. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

8.08 8.00 8.02 8.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 8 degrees angle of attack Figure 126. Comparison of the wing upper surface static pressure distributions at Mach = 0.85 with centerline tail, tw twin canted tails, and tail off; solid LEX.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

12.02 12.09 12.02 12.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (b) 12 degrees angle of attack  Figure 126. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

16.12 16.10 16.03 16.05 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (c) 16 degrees angle of attack  Figure 126. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

20.13 20.06 20.07 20.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (d) 20 degrees angle of attack  Figure 126. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

24.09 24.07 24.05 24.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (e) 24 degrees angle of attack  Figure 126. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

27.98 28.15 28.08 28.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (f) 28 degrees angle of attack  Figure 126. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

12.01 12.06 12.21 12.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 12 degrees angle of attack, -8 degrees angle of sideslip Figure 127. Comparison of the wing upper surface static pressure distributions in sideslip sweeps at Mach = 0.50 wi twin uncanted tail, twin canted tails, and tail off; solid LEX.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

12.07 12.10 12.06 12.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(b) 12 degrees angle of attack, -4 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

12.05 12.05 12.03 12.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(c) 12 degrees angle of attack, +4 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

12.05 11.97 12.15 12.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(d) 12 degrees angle of attack, +8 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

16.08 16.07 16.23 16.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(e) 16 degrees angle of attack, -8 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

16.04 16.06 16.08 16.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(f) 16 degrees angle of attack, -4 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

16.09 16.08 16.10 16.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(g) 16 degrees angle of attack, +4 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

16.05 16.05 16.11 16.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(h) 16 degrees angle of attack, +8 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

20.05 20.03 20.09 20.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(i) 20 degrees angle of attack, -8 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

20.11 20.04 20.08 20.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(j) 20 degrees angle of attack, -4 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

20.12 20.05 20.09 20.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(k) 20 degrees angle of attack, +4 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

20.08 20.07 20.14 20.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(l) 20 degrees angle of attack, +8 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

24.08 24.13 23.54 24.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(m) 24 degrees angle of attack, -8 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

24.11 24.14 23.54 24.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(n) 24 degrees angle of attack, -4 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

24.10 24.10 23.57 24.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(o) 24 degrees angle of attack, +4 degrees angle of sideslip  Figure 127. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

24.08 24.10 23.56 24.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(p) 24 degrees angle of attack, +8 degrees angle of sideslip  Figure 127. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

12.01 12.06 12.21 12.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

 (a) 12 degrees angle of attack, -8 degrees angle of sideslip Figure 128. Comparison of the wing upper surface static pressure distributions in sideslip sweeps at Mach = 0.85 wi twin uncanted tail, twin canted tails, and tail off; solid LEX.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

12.07 12.10 12.06 12.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(b) 12 degrees angle of attack, -4 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

12.05 12.05 12.03 12.01 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(c) 12 degrees angle of attack, +4 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

12.05 11.97 12.15 12.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(d) 12 degrees angle of attack, +8 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

16.08 16.07 16.23 16.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(e) 16 degrees angle of attack, -8 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

16.04 16.06 16.08 16.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(f) 16 degrees angle of attack, -4 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

16.09 16.08 16.10 16.02 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(g) 16 degrees angle of attack, +4 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

16.05 16.05 16.11 16.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(h) 16 degrees angle of attack, +8 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

20.05 20.03 20.09 20.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(i) 20 degrees angle of attack, -8 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

20.11 20.04 20.08 20.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(j) 20 degrees angle of attack, -4 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

20.12 20.05 20.09 20.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(k) 20 degrees angle of attack, +4 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

20.08 20.07 20.14 20.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(l) 20 degrees angle of attack, +8 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

24.08 24.13 23.54 24.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(m) 24 degrees angle of attack, -8 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

24.11 24.14 23.54 24.03 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(n) 24 degrees angle of attack, -4 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

24.10 24.10 23.57 24.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(o) 24 degrees angle of attack, +4 degrees angle of sideslip  Figure 128. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

α, deg 

24.08 24.10 23.56 24.04 

x/c=0.80 

x/c=0.60 

x/c=0.30 

LEX →

(p) 24 degrees angle of attack, +8 degrees angle of sideslip  Figure 128. Concluded.
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

 (a) lift, drag, and pitching moment coefficients 

Figure 129. Comparison of the six-component aerodynamic characteristics at Mach = 0.50 with centerline tail, 

 twin uncanted tails, twin canted tails, and tail off. 

 

 428
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

.05 

.06 

CY 

-.024 

-.020 

-.016 

-.012 

-.008 

-.004 

0 

.004 

.008 

Cn 

-.020 

-.016 

-.012 

-.008 

-.004 

0 

.004 

.008 

.012 

Cl 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

 (b) rolling moment, yawing moment, and side force coefficients 

 Figure 129. Concluded.

 429
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

 (a) lift, drag, and pitching moment coefficients 

Figure 130. Comparison of the six-component aerodynamic characteristics at Mach = 0.85 with centerline tail, 

 twin uncanted tails, twin canted tails, and tail off. 

 

 430
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

CY 

-.016 

-.012 

-.008 

-.004 

0 

.004 

.008 

.012 

.016 

Cn 

-.024 

-.020 

-.016 

-.012 

-.008 

-.004 

0 

.004 

.008 

Cl 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

 (b) rolling moment, yawing moment, and side force coefficients 

 Figure 130. Concluded.

 431
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

0 

.01 

.02 

.03 

.04 

.05 

.06 

.07 

.08 

Cm 

.08 

.09 

.10 

.11 

.12 

.13 

.14 

.15 

.16 

CD 

.44 

.46 

.48 

.50 

.52 

.54 

.56 

.58 

.60 

CL 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

 (a) lift, drag, and pitching moment coefficients at 12 degrees angle of attack 

Figure 131. Comparison of the six-component aerodynamic characteristics in sideslip sweeps at 

 Mach = 0.50 with centerline tail, twin uncanted tails, twin canted tails, and tail off.

 432
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(b) rolling moment, yawing moment, and side force coefficients at 12 degrees angle of attack 

 Figure 131. Continued.

 433
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

0 

.01 

.02 

.03 

.04 

.05 

.06 

.07 

.08 

Cm 

.16 

.17 

.18 

.19 

.20 

.21 

.22 

.23 

.24 

CD 

.56 

.60 

.64 

.68 

.72 

.76 

.80 

.84 

.88 

CL 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(c) lift, drag, and pitching moment coefficients at 16 degrees angle of attack 

 Figure 131. Continued.

 434
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(d) rolling moment, yawing moment, and side force coefficients at 16 degrees angle of attack 

 Figure 131. Continued.

 435
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

.04 

.05 

.06 

.07 

.08 

.09 

.10 

.11 

.12 

Cm 

.24 

.26 

.28 

.30 

.32 

.34 

.36 

.38 

.40 

CD 

.68 

.72 

.76 

.80 

.84 

.88 

.92 

.96 

1.00 

CL 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(e) lift, drag, and pitching moment coefficients at 20 degrees angle of attack 

 Figure 131. Continued.

 436
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(f) rolling moment, yawing moment, and side force coefficients at 20 degrees angle of attack 

 Figure 131. Continued.

 437
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

0 

.02 

.04 

.06 

.08 

.10 

.12 

.14 

.16 

Cm 

.32 

.36 

.40 

.44 

.48 

.52 

.56 

.60 

.64 

CD 

.80 

.85 

.90 

.95 

1.00 

1.05 

1.10 

1.15 

1.20 

CL 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(g) lift, drag, and pitching moment coefficients at 24 degrees angle of attack 

 Figure 131. Continued.

 438
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(h) rolling moment, yawing moment, and side force coefficients at 24 degrees angle of attack 

 Figure 131. Concluded.

 439
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

.05 

.06 

Cm 

.08 

.09 

.10 

.11 

.12 

.13 

.14 

.15 

.16 

CD 

.46 

.48 

.50 

.52 

.54 

.56 

.58 

.60 

.62 

CL 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

 (a) lift, drag, and pitching moment coefficients at 12 degrees angle of attack 

Figure 132. Comparison of the six-component aerodynamic characteristics in sideslip sweeps at 

 Mach = 0.85 with centerline tail, twin uncanted tails, twin canted tails, and tail off.

 440
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(b) rolling moment, yawing moment, and side force coefficients at 12 degrees angle of attack 

 Figure 132. Continued.

 441
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

0 

.01 

.02 

.03 

.04 

.05 

.06 

.07 

.08 

Cm 

.18 

.19 

.20 

.21 

.22 

.23 

.24 

.25 

.26 

CD 

.66 

.68 

.70 

.72 

.74 

.76 

.78 

.80 

.82 

CL 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(c) lift, drag, and pitching moment coefficients at 16 degrees angle of attack 

 Figure 132. Continued.

 442
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(d) rolling moment, yawing moment, and side force coefficients at 16 degrees angle of attack 

 Figure 132. Continued.

 443
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

.02 

.03 

.04 

.05 

.06 

.07 

.08 

.09 

.10 

Cm 

.30 

.31 

.32 

.33 

.34 

.35 

.36 

.37 

.38 

CD 

.76 

.80 

.84 

.88 

.92 

.96 

1.00 

1.04 

1.08 

CL 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(e) lift, drag, and pitching moment coefficients at 20 degrees angle of attack 

 Figure 132. Continued.

 444
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(f) rolling moment, yawing moment, and side force coefficients at 20 degrees angle of attack 

 Figure 132. Continued.

 445
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

.03 

.04 

.05 

.06 

.07 

.08 

.09 

.10 

.11 

Cm 

.40 

.42 

.44 

.46 

.48 

.50 

.52 

.54 

.56 

CD 

.84 

.88 

.92 

.96 

1.00 

1.04 

1.08 

1.12 

1.16 

CL 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(g) lift, drag, and pitching moment coefficients at 24 degrees angle of attack 

 Figure 132. Continued.

 446
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Vertical Tail 

Centerline 

Twin uncanted 

Twin canted 30o

Off 

(h) rolling moment, yawing moment, and side force coefficients at 24 degrees angle of attack 

 Figure 132. Concluded.

 447
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (a) 8 degrees angle of attack Figure 133. Comparison of the wing upper surface static pressure distributions with solid LEX and canard at Mach = 0.5 (Wing-alone data are shown for reference.)
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 10 degrees angle of attack  Figure 133. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (c) 12 degrees angle of attack  Figure 133. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (d) 14 degrees angle of attack  Figure 133. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (e) 16 degrees angle of attack  Figure 133. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (f) 18 degrees angle of attack  Figure 133. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (g) 20 degrees angle of attack  Figure 133. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (h) 22 degrees angle of attack  Figure 133. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (i) 24 degrees angle of attack  Figure 133. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (j) 26 degrees angle of attack  Figure 133. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (k) 28 degrees angle of attack  Figure 133. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (a) 12 degrees angle of attack, -8 degrees angle of sideslip Figure 134. Comparison of the wing upper surface static pressure distributions in sideslip sweeps with the solid LEX and  with centerline tail. (Wing-alone data are shown for reference.)
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 12 degrees angle of attack, -4 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (c) 12 degrees angle of attack, +4 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (d) 12 degrees angle of attack, +8 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (e) 16 degrees angle of attack, -8 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (f) 16 degrees angle of attack, -4 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (g) 16 degrees angle of attack, +4 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (h) 16 degrees angle of attack, +8 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (i) 20 degrees angle of attack, -8 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (j) 20 degrees angle of attack, -4 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (k) 20 degrees angle of attack, +4 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (l) 20 degrees angle of attack, +8 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (m) 24 degrees angle of attack, -8 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (n) 24 degrees angle of attack, -4 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (o) 24 degrees angle of attack, +4 degrees angle of sideslip  Figure 134. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (p) 24 degrees angle of attack, +8 degrees angle of sideslip  Figure 134. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (a) 8 degrees angle of attack Figure 135. Comparison of the wing upper surface static pressure distributions with solid LEX and canard at Mach = 0.8 (Wing-alone data are shown for reference.)
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 10 degrees angle of attack  Figure 135. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (c) 12 degrees angle of attack  Figure 135. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (d) 14 degrees angle of attack  Figure 135. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (e) 16 degrees angle of attack  Figure 135. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (f) 18 degrees angle of attack  Figure 135. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (g) 20 degrees angle of attack  Figure 135. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (h) 22 degrees angle of attack  Figure 135. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (i) 24 degrees angle of attack  Figure 135. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (j) 26 degrees angle of attack  Figure 135. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (k) 28 degrees angle of attack  Figure 135. Concluded.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (a) 12 degrees angle of attack, -8 degrees angle of sideslip Figure 136. Comparison of the wing upper surface static pressure distributions in sideslip sweeps with the solid LEX and  with centerline tail. (Wing-alone data are shown for reference.)
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (b) 12 degrees angle of attack, -4 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (c) 12 degrees angle of attack, +4 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (d) 12 degrees angle of attack, +8 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (e) 16 degrees angle of attack, -8 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (f) 16 degrees angle of attack, -4 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (g) 16 degrees angle of attack, +4 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (h) 16 degrees angle of attack, +8 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (i) 20 degrees angle of attack, -8 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (j) 20 degrees angle of attack, -4 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (k) 20 degrees angle of attack, +4 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (l) 20 degrees angle of attack, +8 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (m) 24 degrees angle of attack, -8 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (n) 24 degrees angle of attack, -4 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (o) 24 degrees angle of attack, +4 degrees angle of sideslip  Figure 136. Continued.
.5 

0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

.5 0 -.5 -1.0 -1.5 -2.0 -2.5 -3.0 

Cp,u 

0 .1 .2 .3 .4 .5 .6 .7 .8 

y/s 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

x/c=0.80 

x/c=0.60 

x/c=0.30 

 (p) 24 degrees angle of attack, +8 degrees angle of sideslip  Figure 136. Concluded.
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

 (a) lift, drag, and pitching moment coefficients 

Figure 137. Comparison of the six-component aerodynamic characteristics with the solid LEX and canard 

 at Mach = 0.50 with centerline tail. (Wing-alone data are shown for reference.) 

 

 502
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

CY 

-.016 

-.012 

-.008 

-.004 

0 

.004 

.008 

.012 

.016 

Cn 

-.008 

-.006 

-.004 

-.002 

0 

.002 

.004 

.006 

.008 

Cl 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

 (b) rolling moment, yawing moment, and side force coefficients 

 Figure 137. Concluded.

 503
-8 -4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

0 .1 .2 .3 .4 .5 .6 .7 .8 .9 1.0 1.1 

CD 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

-.10 -.05 0 .05 .10 .15 .20 

Cm 

-.2 

0 

.2 

.4 

.6 

.8 

1.0 

1.2 

1.4 

1.6 

CL 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

 (a) lift, drag, and pitching moment coefficients 

Figure 138. Comparison of the six-component aerodynamic characteristics with the solid LEX and canard 

 at Mach = 0.85 with centerline tail. (Wing-alone data are shown for reference.) 

 

 504
-4 0 4 8 12 16 20 24 28 32 36 

α, deg 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

CY 

-.016 

-.012 

-.008 

-.004 

0 

.004 

.008 

.012 

.016 

Cn 

-.004 

0 

.004 

.008 

.012 

.016 

.020 

.024 

.028 

Cl 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

 (b) rolling moment, yawing moment, and side force coefficients 

 Figure 138. Concluded.

 505
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

0 

.01 

.02 

.03 

.04 

.05 

.06 

.07 

.08 

Cm 

.08 

.09 

.10 

.11 

.12 

.13 

.14 

.15 

.16 

CD 

.44 

.46 

.48 

.50 

.52 

.54 

.56 

.58 

.60 

CL 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

 (a) lift, drag, and pitching moment coefficients at 12 degrees angle of attack 

Figure 139. Comparison of the six-component aerodynamic characteristics in sideslip sweeps with the solid LEX 

 and canard at Mach = 0.50 with centerline tail. (Wing-alone data are shown for reference.)

 506
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(b) rolling moment, yawing moment, and side force coefficients at 12 degrees angle of attack 

 Figure 139. Continued.

 507
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

0 

.01 

.02 

.03 

.04 

.05 

.06 

.07 

.08 

Cm 

.18 

.19 

.20 

.21 

.22 

.23 

.24 

.25 

.26 

CD 

.64 

.66 

.68 

.70 

.72 

.74 

.76 

.78 

.80 

CL 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(c) lift, drag, and pitching moment coefficients at 16 degrees angle of attack 

 Figure 139. Continued.

 508
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(d) rolling moment, yawing moment, and side force coefficients at 16 degrees angle of attack 

 Figure 139. Continued.

 509
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

0 

.02 

.04 

.06 

.08 

.10 

.12 

.14 

.16 

Cm 

.30 

.31 

.32 

.33 

.34 

.35 

.36 

.37 

.38 

CD 

.76 

.80 

.84 

.88 

.92 

.96 

1.00 

1.04 

1.08 

CL 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(e) lift, drag, and pitching moment coefficients at 20 degrees angle of attack 

 Figure 139. Continued.

 510
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(f) rolling moment, yawing moment, and side force coefficients at 20 degrees angle of attack 

 Figure 139. Continued.

 511
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

0 

.02 

.04 

.06 

.08 

.10 

.12 

.14 

.16 

Cm 

.42 

.44 

.46 

.48 

.50 

.52 

.54 

.56 

.58 

CD 

.92 

.96 

1.00 

1.04 

1.08 

1.12 

1.16 

1.20 

1.24 

CL 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(g) lift, drag, and pitching moment coefficients at 24 degrees angle of attack 

 Figure 139. Continued.

 512
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(h) rolling moment, yawing moment, and side force coefficients at 24 degrees angle of attack 

 Figure 139. Concluded.

 513
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

.05 

.06 

Cm 

.08 

.09 

.10 

.11 

.12 

.13 

.14 

.15 

.16 

CD 

.50 

.52 

.54 

.56 

.58 

.60 

.62 

.64 

.66 

CL 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

 (a) lift, drag, and pitching moment coefficients at 12 degrees angle of attack 

Figure 140. Comparison of the six-component aerodynamic characteristics in sideslip sweeps with the solid LEX 

 and canard at Mach = 0.85 with centerline tail. (Wing-alone data are shown for reference.)

 514
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(b) rolling moment, yawing moment, and side force coefficients at 12 degrees angle of attack 

 Figure 140. Continued.

 515
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

.10 

Cm 

.18 

.19 

.20 

.21 

.22 

.23 

.24 

.25 

.26 

CD 

.66 

.68 

.70 

.72 

.74 

.76 

.78 

.80 

.82 

CL 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(c) lift, drag, and pitching moment coefficients at 16 degrees angle of attack 

 Figure 140. Continued.

 516
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(d) rolling moment, yawing moment, and side force coefficients at 16 degrees angle of attack 

 Figure 140. Continued.

 517
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

.10 

Cm 

.31 

.32 

.33 

.34 

.35 

.36 

.37 

.38 

.39 

CD 

.76 

.80 

.84 

.88 

.92 

.96 

1.00 

1.04 

1.08 

CL 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(e) lift, drag, and pitching moment coefficients at 20 degrees angle of attack 

 Figure 140. Continued.

 518
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(f) rolling moment, yawing moment, and side force coefficients at 20 degrees angle of attack 

 Figure 140. Continued.

 519
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

.10 

Cm 

.40 

.42 

.44 

.46 

.48 

.50 

.52 

.54 

.56 

CD 

.88 

.92 

.96 

1.00 

1.04 

1.08 

1.12 

1.16 

1.20 

CL 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(g) lift, drag, and pitching moment coefficients at 24 degrees angle of attack 

 Figure 140. Continued.

 520
-10 -8 -6 -4 -2 0 2 4 6 8 10 

β, deg 

-.20 

-.15 

-.10 

-.05 

0 

.05 

.10 

.15 

.20 

CY 

-.04 

-.03 

-.02 

-.01 

0 

.01 

.02 

.03 

.04 

Cn 

-.08 

-.06 

-.04 

-.02 

0 

.02 

.04 

.06 

.08 

Cl 

Configuration 

Solid LEX 

Canard 

Wing-Alone 

(h) rolling moment, yawing moment, and side force coefficients at 24 degrees angle of attack 

 Figure 140. Concluded.

 521
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Technical Memorandum

 4. TITLE AND SUBTITLE

Wind Tunnel Investigation of Passive Vortex Control and Vortex-Tail 

Interactions on a Slender Wing at Subsonic and Transonic Speeds 

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

Washington, DC 20546-0001

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 REPORT NUMBER

L-19987

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NASA

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Subject Category 02

Availability: NASA CASI (443) 757-5802

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14. ABSTRACT

A wind tunnel experiment was conducted in the NASA Langley 8-Foot Transonic Pressure Tunnel to determine the effects of passive porosity on vortex flow 

interactions about a slender wing configuration at subsonic and transonic speeds. Flow-through porosity was applied in several arrangements to a leading-edge 

extension, or LEX, mounted to a 65-degree cropped delta wing as a longitudinal instability mitigation technique. Test data were obtained with LEX on and off 

in the presence of a centerline vertical tail and twin, wing-mounted vertical fins to quantify the sensitivity of the aerodynamics to tail placement and 

orientation. A close-coupled canard was tested as an alternative to the LEX as a passive flow control device. Wing upper surface static pressure distributions 

and six-component forces and moments were obtained at Mach numbers of 0.50, 0.85, and 1.20, unit Reynolds number of 2.5 million, angles of attack up to 

approximately 30 degrees, and angles of sideslip to +/-8 degrees. The off-surface flow field was visualized in cross planes on selected configurations using a 

laser vapor screen flow visualization technique. Tunnel-to-tunnel data comparisons and a Reynolds number sensitivity assessment were also performed.

15. SUBJECT TERMS

Aerodynamics; Canards; Flow visualization; Passive porosity; Subsonic speeds; Transonic speeds; Vortex flows; Wind 

tunnels; Wings

18. NUMBER

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 PAGES

530

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01- 04 - 2013