NASA/TP-2001-210629

Flight, Wind-Tunnel, and Computational

Fluid Dynamics Comparison for Cranked

Arrow Wing (F-16XL-1) at Subsonic and

Transonic Speeds

John E. Lamar, Clifford J. Obara, and Bruce D. Fisher

Langley Research Center, Hampton, Virginia

David F. Fisher

Dryden Flight Research Center, Edwards, California

February 2001
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NASA/TP-2001-210629

Flight, Wind-Tunnel, and Computational

Fluid Dynamics Comparison for Cranked

Arrow Wing (F-16XL-1) at Subsonic and

Transonic Speeds

John E. Lamar, Clifford J. Obara, and Bruce D. Fisher

Langley Research Center, Hampton, Virginia

David F. Fisher

Dryden Flight Research Center, Edwards, California

February 2001
Available from:

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Acknowledgments

Many individuals contributed to the gathering of the flight, wind-tunnel, and computational data in this report. A

highly integrated effort was required in order to compare these data. To that end, a by-product of this paper was the

development of a highly skilled and motivated team. This team consisted primarily of direct management and software

support from the Langley Research Center (LaRC) and the Computer Sciences Corporation (CSC), Systems Sciences

Division, Hampton, VA. At LaRC, the key people were Kathryn Stacy, Kurt S. Severance, Catherine K. Cronin,

Donald P. Randall, Patricia A. Kerr, Michael D. Weisenborn, Susan J. Rickard, and David E. Hahne; at CSC, the key

people were Christine G. Matthews, Dana P. Hammond, Raymond L. Gates, Robert C. Auberg, Joseph F. Reisel,

Gerald H. Mall, Susan S. Bowen, S. Blake Williams, Rachel Gartenberg, Norma F. Bean, Javier A. Garriz, and

David A. Wolverton. Individual contributions ranged from data translation done by Patricia A. Kerr, Dana P.

Hammond, and Robert C. Auberg; image fusions done by Kathryn Stacy; special data comparisons done by Kurt S.

Severance, David A. Wolverton, Gerald H. Mall, Susan S. Bowen, S. Blake Williams, Rachel Gartenberg, and

Stephen J. Alter, Lockheed Martin Engineering and Sciences Corporation, Hampton, VA; data base management work

done by Catherine K. Cronin, Donald P. Randall, Joseph F. Reisel, and Raymond L. Gates; initial geometry, computational gridding, and solution work done by Anutosh Moitra, High Technology Corporation, Hampton, VA; final

geometry and computational gridding work done by Norma F. Bean; initial computational solutions on final grid performed by Javier A. Garriz; computational camera location and viewing work done by Michael D. Weisenborn; and

low-speed wind-tunnel data provided by Susan J. Rickard and David E. Hahne.

The authors wish to thank the following people from LaRC for the many helpful suggestions arising from their review

of this manuscript: Laurence D. Leavitt, Jerry M. Allen, Scott G. Anders, Jay M. Brandon, Wendy B. Lessard, Kurt S.

Severance, and Georgina R. Branco of NASA Dryden Flight Research Center. Also, thanks are given to the Library

and Media Services Branch of the Langley Research Center who published this document, with special

acknowledgment to the NCI Technical Publications Support Group at Langley (Leanna Dee Bullock, Eloise Johnson,

Una Thomas, Martha Jones, Patricia Gottschall, Mary Edwards, and Cathy Everett).

The use of trademarks or names of manufacturers in this report is for accurate reporting and does not constitute an

official endorsement, either expressed or implied, of such products or manufacturers by the National Aeronautics and

Space Administration.
Contents

Symbols . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . v

Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

Geometry ....................................... 1

Flight, CFD, and Wind-Tunnel Surface Pressures . . . ................... 1

Boundary Layer . ................................... 1

Data Fusion . . . ................................... 1

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

Basic Data Types and Test Conditions ........................... 2

Geometry ....................................... 2

Airplane description ................................ 2

Details and comparisons . . . ........................... 3

Flight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3

Data collected ................................... 3

Procedures and accuracy of data collection . . . ................... 3

Pressures . . ................................... 4

Visual . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

Hot lm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5

0.11-scaled model tests ............................... 5

0.18-scaled model test ................................ 5

0.04-scaled model test ................................ 6

CFD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

CFL3D 
ow solver code and grid ........................... 6

Computer requirements and sample convergence histories ................ 6

Using the solution le ................................ 6

CFD Comparisons . ................................... 7

Surface Pressures . ................................... 7

Subsonic . . ................................... 7

Transonic . . ................................... 8

Contours . . ................................... 9

Boundary-Layer Proles at FC 7 . . ........................... 9

Skin Friction at FC 7 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

On and O Surface Flow Features at FC 46 . . . . . . . . . . . . . . . . . . . . . . . . 11

EFD Basic Data and Comparisons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

Forces and Moment in Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

Surface Pressures in Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11

Eect of model scale and Rn on Cp at low speed . . . . . . . . . . . . . . . . . . . . 11

Transonic Cp;u contours . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

iii
On and O Surface Flow (Air Dams O) in Wind Tunnel . . . . . . . . . . . . . . . . . . 12

Surface Flow in Fli ght . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

Concluding Remarks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

Appendix A|Visualization Tools . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

Appendix B|Data Processes and Management . . . . . . . . . . . . . . . . . . . . . . . 17

Appendix C|Airplane Instrumentation . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

Appendix D|Description of Wind-Tunnel Test and Vapor-Screen Systems . . . . . . . . . . . 29

References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33

Tables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35

Figures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63

iv
Symbols

A wing aspect ratio, b2

Sref , 1.75

AMRAAM advance d medium range air-to-air missile

BART Langley Basic Aerodynamic Research Tunnel

BL butt line on airplane , positive on right wing, in. (see g. 2)

b wingspan, 32.4 ft (full scale)

CD drag coecie nt

CL lift coecient

Cm pitching-moment coecie nt

Cp static-pre ssure coecie nt

CAWAP Cranked-Arrow Wing Aerodynamics Pro ject

CFD computational 
uid dynamics

cf local skin friction coecient

cr the ore tical root chord for F-16XL airplane , 41.75 ft (full scale)

c re ference wing chord, 24.7 ft (full scale)

EFD experime ntal 
uid dynamics, 
ight- or ground-based measurements

ESP ele ctronic scanning pressure

FAST Flow Analysis Software Toolkit

FC 
ight condition

Flt 
ight number

FS fuselage station on airplane, positive aft, in. (se e g. 2)

FV 
ow visualization

GPS global positioning system

HF hot-lm se nsors

HUD he ads-up display

h altitude , ft

I.D. inside diame te r

i; j; k grid indice s

L2 (re sidual) re sidual of the L2 norm plotted on log scale

LE leading e dge

LS, low lowe r surface

LT loaded turn

M Mach number

M1 fre e-stre am Mach number

NAS Numerical Aerodynamics Simulation

O.D. outside diameter

PGME propyle ne glycol methyl ether

v
PrISM Program for Information Storage and Management

PSP pressure sensitive paint

PT modied Preston tubes

p free-stream static pressure, lb/ft2

q1 free-stream dynamic pressure, lb/ft2

Rn Reynolds number, 
ight value computed by 1:2583cpM1

 T + 198:72

T2



106 (based on

ref. 1, eq. I.3.8-(1)); in gure keys, R is used for Reynolds number

R Reynolds number based on momentum thickness as computed by  0:407 sin (70) h Rn=ft

cos (70)i0:5

RK boundary-layer rake

r radius

Sre f reference wing area, 600 ft2 (full scale)

SLFC supersonic laminar 
ow control

T absolute temperature, R

TE trailing edge

US,up upper surface

V=VRE ratio of velocity magnitude in boundary layer to that at rake extreme total-pressure

tube

VCR video cassette recorder

WL waterline on airplane, positive up, in. (see g. 2)

WT wind tunnel

WUT wind-up turn

WWW World Wide Web

x=c fractional distance along local chord, positive aft

y normal distance above surface at rake location, in.

y+ Reynolds number like term for 
at-plate turbulent boundary layer (ref. 2)

 angle of attack, deg

 angle of sideslip, deg

; ; ! Euler angles for camera orientation, deg (see g. C6)

2y=b fractional distance along local semispan, positive on right wing

2-D two-dimensional

30 by 60 Langley 30- by 60-Foot Tunnel

3-D three-dimensional

Subscripts:

avg average value

nom nominal value

u upper surface

vi
Summary

This report summarizes an eort to compare

geometrical, 
ight, computational 
uid dynamics

(CFD), and various wind-tunnel studies for the

F-16XL-1 airplane and uses data fusion (overlaying) wherever possible to provide additional insight

into the data sets. The resulting highly diverse

types of data sets were obtained over a wide range

of 
ow test conditions and have produced some

novel results.

Geometry

With respect to geometry, the upper surface of

the airplane and the numerical surface description

have a reasonably good comparison. The largest

dierence measured over the wing was 0.24 in.

Flight, CFD, and Wind-Tunnel Surfa ce

Pressures

For 
ight and CFD pressures, overall comparisons have been expedited by use of computational

graphical tools. At subsonic speeds, reasonably

good agreement was noted at angles of attack up to

10.4. For angles of attack above 10.4, the agreements were much poorer because the suction peaks

were underpredicted. At transonic speeds, generally good pressure agreement was noted, especially

near the inboard leading edge at angle of attack

() of 4.4 and free-stream Mach number (M1)

of 0.97. The eect of small trailing-edge controlsurface de
ections ( 2) in 
ight, not modeled in

CFD, led to poorer agreement for the aft pressures

and a shock not predicted. By contrast, an experimental 
uid dynamics comparison of the pressure contours from the wind tunnel at  = 4 and

M1 = 0:95 with the aforementioned 
ight data did

show a wind-tunnel shock location near the hinge

line for the aileron, which generally is in qualitative

agreement with that determined from the transonic


ight results.

Boundary Layer

Anticipated relaminarization of the boundary

layer under the in
uence of high suction peaks near

the leading edge was not supported by the turbulent signal output from 
ight hot-lm gauges.

Thus, the 
ight leading-edge attachment line was

turbulent; this puts this data set in agreement with

other 
ight data. The grid used in the CFL3D code

produced Reynolds number like term for 
at-plate

turbulent boundary layer (y+) values much larger

than 2 at 
ight conditions; in fact, they were 82.

The \wall function" option in the code proved su-

cient at subsonic speeds to represent the boundary

layers underneath attached 
ow and the primary

vortex but insucient underneath the secondary

vortex system at nominal angle of attack of 13.

Moreover, the 
ight data indicate the presence of

an interaction of the secondary vortex system and

boundary layer, and the boundary-layer measurements show that the secondary vortex was located

more outboard than predicted. Qualitative agreement exists between predicted and measured skin

friction spanwise distributions because both CFD

and 
ight results show at least two regions of high

local skin friction coecient (cf), indicative of primary and secondary vortex presence, and they occur at spanwise positions near one another. However, the absolute levels of cf and the indicated

spanwise extent of the vortices dier.

Data Fusion

Data fusion (overlaying) has been applied to a

variety of 
ow-visualization techniques for 
ight,

wind-tunnel, and CFD investigations with insightful results. Among these are (1) combined surface

tufts images from three 
ight cameras compared

with CFD surface streamlines and vortex-core representation at M1 = 0:53 and = 10.4 show generally good comparison; (2) the 
ight results of oil


ow with tufts and liquid crystals with tufts compared well; and (3) the wind-tunnel comparison

of vortex-core and reattachment-point locations|

as determined from vapor screen and

compared with surface oil-
ow traces|though

showing promise, point to the need for further re-

nement.

Introduction

The increasing capability of CFD to simulate

real airplane geometries and 
ow conditions is

found in the literature for both ghter and transport airplanes, for example, the F-18 HARV (ref. 3)

and the MD-11 (ref. 4). In the continuance of that

trend, this report details the results of a comparative study of CFD, 
ight, and wind-tunnel data for

the F-16XL-1 airplane (g. 1) over a wide range of

test conditions from subsonic to transonic speeds.

Though this airplane is not new, its cranked-arrow

planform is relevant to any high-speed (supersonic)

ghter or transport, for example, the High-Speed
Civil Transport. When the F-16XL-1 airplane

was new, the kinds of testing done were related

to its operational characteristics (refs. 5 and 6)

and not focused on an understanding of its basic 
ow physics, the sub ject of the current investigation. This understanding was the subject of

the Cranked-Arrow Wing Aerodynamics Project

(CAWAP) and the results reported herein. Since

CFD predictions are being sought for purposes of

comparison with measured data for this airplane,

knowledge of the airplane geometry is rst necessary. Hence, a geometrical comparison was undertaken for the complete airplane, including wingtipmounted missiles and tail. Other comparisons are

made between 
ight, wind-tunnel, and CFD data.

The scope of the planned comparisons was extensive, as shown in table 1; however, not all the

needed data were collected during the testing process or proved reliable with the latter essentially

eliminating the load-factor and Rn 
ight studies.

Along with other constraints, this has led to the

availability of only a subset of comparisons as detailed in table 2. The reasons for the dearth of

data collection range from pressure sensitive paint

(PSP) not yet calibrated for use in 
ight, along

with other imaging and processing issues; hardware

for the 
ight vapor screen not available during the

testing time; and resource constraints forced the

elimination of the propylene-glycol-methyl-ether

(PGME) trace 
ight experiment. In addition to

these, at least two other planned tasks were also

not fully accomplished. For example, (1) the modied Preston tubes used to determine local skin

friction were to be located at the same fuselage

station as the boundary-layer rakes, but a more

aft position for the tubes was necessitated because

of easier installation on the airplane and avoiding

the 
ow o a step in the leading-edge region and

(2) an attempt to correct the transition point location used in CFD from the leading edge to some

point aft proved unnecessary at 
ight conditions

because the leading edge was already turbulent according to hot-lm measurements.

Notwithstanding these issues, a number of comparisons were made; these include on and o surface 
ow characteristics, that is, surface pressures,

upper surface 
ow, boundary-layer velocity pro-

les, and local skin friction distribution, as well as

force and moment where possible. These comparisons are followed by analyses of data sets within a

given type, for example, wind tunnel only. (Ref. 7

details what was learned about the boundary-layer


ow over the F-16XL-2 airplane|the one used in

the supersonic laminar 
ow control (SLFC) 
ight

experiments.)

In addition to the basic comparison eort, a

brief introduction on the use of associated visualization software tools is given in appendix A, and

a description of the data processes and mangement

system used for this project is given in appendix B. Appendixes C and D contain a description

of the airplane instrumentation used and the associated wind-tunnel test and vapor-screen systems,

respectively.

Basic Data Types and Test Conditions

Four basic data types are associated with this

comparison eort. Each type is important in its

own right and when utilized together form the

basis for this comprehensive undertaking. The

data types are geometry, 
ight, wind tunnel, and

CFD.

Geometry

Airplane description. The F-16XL-1 is a

single-place ghter-type prototype airplane developed by the General Dynamics Corporation,

Ft. Worth Division (now the Lockheed MartinTactical Aircraft Systems) by stretching the fuselage of a full-scale development F-16A and adding

a cranked-arrow wing, a modied fuel system, and

a modied 
ight control system. This airplane had

scheduled leading-edge 
aps, elevons, and ailerons

on the wing for control. The technical specications for the airplane are given in table 3. Details

on the construction of the airplane and its intended

missions are given in references 5, 6, and 8.

The design of the cranked-arrow wing was a

cooperative eort of the Langley Research Center and the General Dynamics Corporation. The

new wing was designed to provide the F-16 airplane

with improved supersonic performance while maintaining transonic performance comparable with

that provided by the current F-16 design. As

shown in gure 2, the resultant design had a

leading-edge sweep angle of 70 inboard and 50

outboard of the crank. At the juncture of the wing

leading edge with the fuselage, an \S-blend curve"

was placed in the leading edge to alleviate a pitch

2
instability that occurred at high angles of attack

in wind-tunnel tests. Because the wing sweep and

general arrangement of the cranked-arrow wing of

the F-16XL-11 resembled the planform of a proposed high-speed commercial transport concept

(g. 3), this was reasoned to be a sucient basis for using the airplane as a platform on which

to conduct high-lift research for the NASA High

Speed Research Program. All 
ight tests reported herein were with the air dams|upper surface fences mounted near the wing leading-edge

crank|and wingtip missiles installed, as shown in

gures 1(c) and 2.

Details and comparisons. Much eort was expended to have a consistent airplane, wind-tunnel

model, and computational geometry so that comparisons between the various data sets of measured

and computational results would be meaningful.

To facilitate this consistency assessment, the measured geometries were compared with the numerical surface as generated from the geometry reports

(refs. 9 and 10), les of lofting details, and digitization of a mold made from a portion of the airplane

inlet region. This assessment was only fully done

for the 
ight vehicle.

A partial geometrical assessment of the 0.04-

scaled model determined that the wing panels had

a slight asymmetry about the centerline. This

model had no force/moment balance; therefore, the

data collected for use herein are the oil 
ow, vapor

screen, and surface pressures. Because the oil-


ow and vapor-screen results are compared on the

same wing with no other data, they are consistent.

Moreover, PSP results obtained on both wings

(ref. 11) show only a slight eect of the geometrical

asymmetry on the results.

On the airplane, the upper surface of the geometry has been quantied to within 0.002 in.

by using photogrammetry and 1089 targets. (See

g. 4.) The photogrammetry was done with the

airplane defueled (only unpumpable fuel remaining), leveled on jacks, and the control surfaces set

to zero. Figure 5 shows both the measured ge1 The airplane was on loan to NASA from the U.S. Air Force

F-16 Spe cial Projects Oce at Wright-Patterson Air Forc e Base,

Ohio. The loan agreement permitted NASA to conduc t re -

search 
ights at both Dryden Flight Research Center and Langley

Research Ce nter.

ometry and the numerical surface for the overall conguration (g. 5(a)) and detailed cross sections (gs. 5(b) to (e)). The overall agreement is

seen to be reasonable, with the maximum dierence over the wing being 0.24 in. However, over

the airplane forebody length of 92 in., gure 5(e)

shows the maximum upper surface dierences to

be larger, reaching a vertical value of 1.65 in. or a

normal-to-surface value of 0.91 in. near the nose|

the measured surface was below the numerical one.

This dierence leads to a nose droop angle of only

1. An implied assumption here is that the forebody lower surface follows the upper surface in

the de
ection, which is reasonable, and since no

lower surface photogrammetry measurements were

made, it is a logical choice. This value of nose

droop angle should have minimal impact on the

pressure results but could slightly impact lift at

 = 0 and pitching moment at zero lift. The relatively large dierence at the nose, in comparison

with those on the wing, is not fully explainable

but is in the direction of structural de
ection associated with cantilevered weights mounted ahead

of the windshield. The weight is in two forms|

ballast (674 lb) and racks of instruments.

Flight

Data collected. Seven dierent kinds of 
ight

data were collected; four are shown schematically

in gure 6. Three were pressure based|surface

static pressures, boundary-layer rakes, and modied Preston tubes (ref. 12); three were video

recorded|surface tufts, surface oil, and surface

liquid crystals; and one was hot-lm data. The

pressure and surface-
ow data are used for the

purpose of establishing the eects of variation in

Mach number on the local 
ow. These data serve

as the basis for comparison with other data sets.

The hot-lm data are used to establish whether

boundary-layer transition occurs and under what

test conditions.

Procedures and accuracy of data collection.

The subsonic 
ight data were obtained at stabilized 
ight conditions at 5000, 10 000, 17 500, and

24 000 ft at nom = 5, 10, 13, 15, and 20 at

 = 0 for 30 sec. Each stabilized point was followed by a slow  sweep and held at  5 for at

least 5 sec. Subsonic data were also obtained in stabilized level turns at altitudes of 5000, 10 000, and

3
17 500 ft at M1 = 0:51 and  = 10, at M1 = 0:42

and  = 13, and at M1 = 0:38 and  = 15.

By varying the altitude, transonic and supersonic

data were obtained at M1 = 0.9 to 1.5 and constant Rn = 3:6 106/ft at stabilized 1g conditions

and in slow wind-up turns (WUTs). At M1 = 1:3,

the data were obtained in stabilized 1g 
ight conditions and slow WUTs at Rn from 2:0 106/ft to

5:0 106/ft.

For M1, the accuracy of key 
ight parameters

is 0:003 at subsonic speeds and 0:005 at supersonic speeds; for , 0:3; for , 0:5; and

for surface static pressure, 0:024 psi for the 5-psi

module and 0:048 psi for the 10-psi module.

Pressures. The right-wing surface pressures|

mostly upper surface|were measured by using

326 static ports, both 
ush and in streamwise

belts, through eleven 32-port electronically scanning pressure (ESP) transducers, also called modules. (For the leading-edge ports, 10-psi modules

were used and elsewhere 5-psi modules were used.)

Table 4 shows that 360 ports were planned but

only 337 were plumbed because of aircraft structural constraints or other valid reasons; of these

ports, only 280 on the upper surface and 46 on the

lower surface proved reliable. The ports were arranged so that a sucient number would be at a

given BL or FS for cross plotting, as well as for covering other regions of special interest, that is, the

apex and ahead of and behind the hinge lines of

the trailing-edge control surfaces. Figure 7 shows

a schematic layout of the requested arrangement

based on the port pattern of wind-tunnel model

and the actual or as-
own arrangement on the airplane. Table 4 contains the requested, actual, and

nominal port locations. Nominal FS and BL values

for the ports were needed in order to facilitate comparisons between the dierent data sets; moreover,

adjusting the 
ight BL values was needed in order

to make chordwise plots of Cp. This need is due to

the BL locations of the streamwise ports diering

even on the same streamwise belt of pressure tubes

because the belt was composed of a group of tubes

located nominally at the requested location.

Figure 8 details the complete planned pressure instrumentation system layout on the airplane, including the distribution of the static ports

by type, belt or 
ush, and boundary-layer rakes

or modied Preston tubes. These static ports

are connected to internally mounted ESP modules through 0.0625-in-O.D. tubes (0.028-in. I.D.).

Each pressure tube in the belt was used to measure two separate values of pressure. This measurement was accomplished by sealing each tube

about halfway along its length; thereby provision

was made for one forward and one aft port. The

numbers associated with the belt static ports in

gure 8 were the values of the actual ports for that

belt. Figure 9 shows a photograph of the pressure

instrumentation suite 
own on the right wing.

Boundary-layer measurements were made by

using two rakes at a time at four dierent positions on the left wing with the most inboard one

always used as a control. Each rake used 16 active tubes, 15 total pressure and 1 static pressure, of the 23 available. These two rakes were

connected to one 32-port ESP module located inside the left wing. Figure 10 shows a sketch of

the rake (overall height 2 in.) and denotes the active tubes. When mounted on the airplane, each

rake was oriented into the local 
ow at an average angle over its height based on initial CFD predictions from the CFL3D code (refs. 13 and 14).

The 
ow conditions were for the complete airplane

(half-airplane modeled with symmetry assumed) at

 = 13, M1 = 0:29, and Rn = 46:1106, that is,


ight condition (FC) 7 (table 5). Figure 11 shows

the four locations chosen|one well inboard of the

shed vortex systems, one underneath the primary

vortex, and two associated with the secondary vortex, both underneath and at its separation point;

all are at a nominal position of FS 295 along the

predicted orientation which takes into account the


ow at and slightly o the surface. The average

of these local 
ow directions was used to establish the rake orientation angles. These angles were

measured from the centerline with the rake pointing forward and inboard and have values of 7.5,

45, 27.5, and 23.5, for rakes 3, 4, 7, and 5,

respectively.

The 16 modied Preston tubes (g. 6)|the

modication to each Preston tube is the integration of a static-pressure port with the totalpressure tube|are used for the determination of

local skin friction across the left wing near FS 330.

(See ref. 15 for a discussion of the basic Preston

tube.) These 32 pressures use the same ESP as

the rakes but not on the same 
ight. The tubes

were aligned with the local 
ow by using the same

4
initial CFL3D solution at FC 7. The equation

used to generate the experimental 
uid dynamics

(EFD) cf values comes from reference 12 and re -

lates, through a process of calibration, the pressure

change between the total- and static-pressure tubes

to the local skin friction.

Visual. Video data were recorded with up to

six external cameras: two mounted atop the vertical tail, one on either side of the fuselage behind

the canopy, and one in the nose of each dummy

missile. An internally mounted heads-up display

(HUD) camera was also used on occasion. Figure 12 shows the camera locations on the airplane.

The time was added to each image by a time-code

inserter (appendix C) so that the images could

be compared to form a composite and the 
ight

test conditions could be established. (See table 6

for the video-image times for the nominal 
ight

conditions.) Images of interest were digitized in

a 512- by 480-pixel format for further processing

to develop quantiable video data. In addition

to the images, the other input quantities needed

for the processes, given in appendix A, are the

video targets (shown schematically in g. 7) and

the position and calibration characteristics of each

camera/lens combination.

Hot lm. Table 5 shows the minimum value

of R to be 122. Reference 16 indicates that a

laminar 
ow reattachment line is not expected for

this 3-D wing unless R is less than 100. Even

though the calculated values of R for this test exceed those expected for laminar 
ow, 
ow relaminarization was anticipated to occur because of the

high suction peaks in the vicinity of the leading

edge. Figure 13 shows the hot-lm belt installed

on the airplane. Although there were 24 gauges in

a belt, only 12 were active on any one 
ight.

The anticipated relaminarization of the boundary layer along the leading edge for the F-16XL-1

was not evident in the traces made from these

hot-lm gauges; therefore, only turbulent 
ow was

present over the airplane.

The F-16XL-2 airplane was used for SLFC 
ight

experiments and employed hot-lm sensors, as described in reference 7. The dierence in application

of the sensors was in the placement. For CAWAP,

the sensors were around the leading edge in a belt,

whereas for the SLFC, they were at various chord

locations.

Wind Tunnel

Table 7 provides a summary of the data previously collected as well as that taken specically for

this research project. Note the variety of model

scales, Mach number, and Reynolds number over

which tests were conducted. As can be seen from

the table, not all models had the air dams installed;

this prevented those data sets from being utilized in the ensuing comparisons. In general, the

vast majority of the comparisons presented are for

data sets from complete congurations; however,

some Langley Basic Aerodynamic Research Tunnel (BART) results for a conguration subset|

air dams o|are shown in the section \EFD Basic Data and Comparisons."All wind-tunnel data

are for the conguration with the control surfaces

unde
ected.

0.11-scaled model tests. The tests for the

0.11-scaled model tests were conducted in the

Ames 11-Foot Tunnel on the model described in

reference 17 and with the tabular results reported

in reference 18. This model was specically built

to estimate the airloads for the airplane from

M1 = 0:60 to 2:0 using 109 ports on the left wing

upper surface and 81 on the right wing lower surface. The ports are located in streamwise rows and

the positions are given in table 8 at airplane scale

for both actual and nominal values. None of these

streamwise rows nominally match those chosen for

the airplane. Because of data release restrictions

of these wind-tunnel Cp results, no direct comparisons are made with 
ight or CFD; however, unscaled transonic Cp;u contours are presented later.

0.18-scaled model test. The test for the 0.18-

scaled model was conducted in the Langley 30-

by 60-Foot Tunnel; the pressure results have

not been formally published elsewhere, but the

force/moment results have been published in reference 19. The basic model has 30 right-wing, 
ush

upper surface ports located to yield pressures for

both streamwise and spanwise rows. These ports

are given in table 9 at airplane scale for both actual

and nominal values. These locations do not duplicate those on the 0.11-scaled model. The ports

on the 0.18-scaled model are nominally duplicated

5
on the airplane. Only two runs|one force/moment

and one pressure|are utilized from this test, and

both are at  = 0.

0.04-scaled model test. The test for the 0.04-

scaled model was conducted in the BART at

 = 0. This model has 82 ports, 57 on the rightwing upper surface and 25 on the left-wing lower

surface; the actual and nominal values are tabulated in table 10 at airplane scale. These 82 ports

were specied to encompass the 30 ports of the

0.18-scaled model used in the 30- by 60-Foot Tunnel test. Moreover, the ports on the 0.04-scaled

model are nominally duplicated on the airplane.

Selected results from the BART test appear in reference 20, where they have been compared with

Euler code predictions.

CFD

CFL3D 
ow solver code and grid. The 
ow

solver code CFL3D was run in the Navier-Stokes

mode with a turbulent boundary layer employing

the Baldwin-Lomax with the Degani-Schi turbulence model (in the j-k directions) on a multiblock,

patched grid over a variety of wind-tunnel and


ight test conditions. Two separate grids were used

to model half the airplane conguration (with unde
ected control surfaces) and external 
ow eld.

The initial grid had 36 blocks and was used with

version 3 of the 
ow solver to produce the initial

results upon which the locations and orientations

of the surface instrumentation were set. The current grid had 30 blocks and was used with version 5 of the 
ow solver to obtain the comparative

solutions reported herein. The current grid was

needed for two reasons: (1) to have the grid more

closely conform to the actual fuselage and wing geometries and (2) to improve the grid layout on the

wing and fuselage surfaces. Figures 14 and 15 provide the two grids on the wing-fuselage surface and

outer boundary, excluding the missile and missile

rail grids, which were eectively unchanged. For

the current grid, the inner region of the airplane

was modeled by 16 blocks, the outer region by 14

blocks, and all 30 blocks are shown schematically

in gure 16. The boundary conditions were symmetry, solid wall for the outer mold lines, 
ow into

the duct inlet with the exhaust face faired over, and

Riemann-type conditions at the far-eld boundaries. A total of 1 372 096 cells (1 707 117 node

points) were used to obtain solutions at specied

test conditions (i.e., , M, grid Rn, T ). To maximize computer resource allotments, the minimum

number of cells was used. The resulting grid spacing normal to the numerical surface led to a value

of y+ of 2 at wind tunnel Rn, whereas at 
ight

Rn the average value of y+ was 82. In an eort

to compensate for the insucient grid spacing at


ight conditions, the \wall function" option was

used to augment the turbulence model in CFL3D.

The wall function is dened as that boundary-layer

growth rate expected from a turbulent mean 
ow

near the wall (ref. 21).

Thirteen separate CFD solutions were obtained

with the current grid and version 5 of CFL3D, and

they are referenced in the Program for Information

Storage and Management (PrISM) data base which

is described in appendix B. In PrISM, these solutions are numbered from CFD:CFL3D(001)Run:1

through CFD:CFL3D(001)Run:13, with CFD:

CFL3D(001)Run:4 and CFD:CFL3D(001)Run:12

not being converged, as shown in table 11. The

latter group of twelve had improvements made in

the grid and/or boundary-layer modeling, and selected ones from this set are used in the comparisons presented herein.

Compu ter requirements and sample convergence histories. All computations were performed on an NAS Cray C90, located at the Ames

Research Center. On this machine, the algorithm

uses 65 megawords of storage and requires approximately 13 sec/cell/cycle. Converged results were

nominally obtained in 5850 cycles requiring about

24 hr of computer time for the grid. (The minimum and maximum values of cycles/computing

time were 5350/20 hr and 16 850/79 hr, respectively.) The number of cycles was sucient to reduce the residuals by 2 to 3 orders of magnitude

and to reduce the oscillations in CL and CD to a

negligible level. The computations were performed

with multigrid and mesh sequencing. Two examples are shown in gure 17, one at high Rn and one

at low Rn. The high Rn solution for FC 49 shows

good convergence properties, whereas the low Rn

solution for the BART test shows some small oscillations both in terms of L2 (residual) and CL

through 10 400 cycles.

Using the solu tion le. The solution le

was readily usable in a standard postprocessing

6
package, as described in appendix B for surface

pressure plotting; however, the processing needed

to determine the velocity magnitudes for comparison with the boundary-layer rake data was not as

straightforward. There were two reasons: (1) the

initial and current grids diered over the surface

so that the locations of the rakes did not occur at

current grid points and (2) the velocities needed to

be established along a normal to the surface at the

specied points in order to be comparable with the

rake data. Since most o the surface grid lines were

not normal to the surface over the length needed, it

was necessary to have software written to perform

the data interpolation|the one chosen was IBM

Data Explorer software (ref. 22)|and to establish

these velocities along surface normals at the rake

locations.

The cf distribution, calculated in CFL3D on

the wing upper surface near FS 330, has been

extracted from the results les for the appropriate

inboard, upper wing surface blocks and is placed

into a single le for later comparison with the EFD

results.

CFD Comparisons

This section contains CFD comparisons with

measured surface pressures, boundary-layer pro-

les, local skin friction, and on and o surface 
ow

features.

Surface Pressures

Comparisons of surface pressures were expedited by using the World Wide Web (WWW) for

accessing the data les on the Langley mass storage

system, as explained in appendix B. Flight pressure

data are primarily available for 
ights 143 to 154

which had good image data and are denoted in table 5 by a lled cell. However, it turned out that

not all the runs with good images in these 
ights

had pressure data. To distinguish between the two,

the table shows the ones with pressure data colored yellow and the ones missing the data colored

red. Regardless of the image quality for the runs

in 
ight 145, its pressure calibrations were not resolvable and hence the data not recoverable. (Note

that the nominal values of test parameters are used

only for identifying similar data sets for comparison, but in the gure caption the actual values of

test parameters are given.)

Because of changes in Rn, dierences between

the 
ight and wind-tunnel pressure data sets are

expected to range from slight to large for most

test parameters. Additional slight changes are

due to small dierences in  and M1 at a nominal test point. Larger dierences are associated with geometrical dierences of control-surface

(leading-edge 
ap, elevon, and aileron) de
ections and aeroelastic deformations on the outboard

wing panels during 
ight. In addition, there may

be some 
ight pressure-eld distortion caused by

the 
ow over the pressure belts. (The preferred

method of measuring surface pressures is with 
ush

orices, as done in the wind-tunnel tests and in

the airplane leading-edge region; however, with

the wing containing fuel|precluding the penetration of the skin for such orices|pressure belts

were used to provide the bulk of the 
ight surfacepressure measurements.) All 
ight results presented herein are at 1g conditions. For reference

purposes, the white line on the airplane or model

indicates the location of the row of Cp ports.

Subsonic. Eight separate CFD subsonic solutions were obtained (control surfaces unde
ected)

of which seven simulated 
ight test conditions and

one simulated a BART test condition. The 
ight

solutions correspond to FC 1, 7, 19, 25, 34, 46, and

49 from table 5. Unfortunately, the surface pressure data for FC 7 and 19 are not available; hence

only data for the other ve FCs are oered in gures 18, 19, 20, and 21 at nom = 5, 10, 13, and

20, respectively. The fact that 
ight pressures are

missing for FC 7 is especially noteworthy because

they were for one of the chief 
ight conditions at

nom = 13 and Mnom = 0:29.

Three general observations can be made from

the cited group of gures: the lower surface Cp

data are well predicted; the upper surface Cp data

are generally well predicted inboard of the crank;

and all 
ight Cp data show an unusual and uncharacteristic variation near 2y=b = 0:6 for FS 185.

The ports for the latter comment are located in and

towards the aft-end portion of the S-blend curve

part of the airplane, a region of the airplane where

the 
ow is very sensitive. Even with this unusual

behavior of Cp, these results are retained because

they do bracket the predicted data at these subsonic values of M1. Figures 18 and 19, at the lower

values of  of 5.5 and 10.4, show the main Cp

7
features, including suction peaks, to be generally

well predicted, both chordwise and spanwise, but

not in all details. Whereas, at both higher values

of  of 13.0 and 20.0, gures 20 and 21 show the

suction peaks to be underpredicted; the spanwise

distributions are slightly better predicted than the

chordwise but neither one really well.

Figures 18(o) and 19(o) present overall comparisons of CFD and 
ight Cp data for FC 1 ( = 5:5)

and FC 46 ( = 10.4), respectively. These gures were developed by using FAST (ref. 23) with

the CFD solution data serving as the background

color mirrored about the centerline and with the


ight Cp data superimposed. (These data are in

PLOT3D (ref. 24) format.) The 
ight port locations are denoted as black dots and the associated

Cp values by the color of the surrounding bubble outline. Both gures show the general overall

agreement to be good in that the colors of the bubble outlines are indistinguishable from the CFD

surface, with its associated Cp color bar, over a

large part of the wing. More dierences appear to

occur between the data sets in gure 18(o) than

in gure 19(o), because gure 19(o) has a more

restricted Cp range. Where dierences are noted

in gure 18(o), the 
ight values are seen to be

more positive than the CFD values with the exception of six ports in the apex region. Figure 19(o)

shows some similar patterns; where dierences do

occur, the measured values are generally more positive. However, they are primarily restricted to

the forward part of the wing. The exceptions occur in the apex region where the insert illustrates

the measured suction pressures for ve ports to be

more negative than predicted, that is, a region of

strong vortex in
uence. Figures 18 and 19 show

the need for both the detail and overall Cp comparison graphs.

Figure 21 also contains the BART Cp data and

CFD predictions. A comparison shows the agreement of the BART data with predictions to be

no better than that noted for 
ight. However,

there are oscillations in the predicted BART Cp

observed over the aft part of the wing, which may

be associated with the oscillations noted in the

solution (g. 17(b)); there is nothing in the 
ow

physics to suggest such an oscillatory Cp behavior. This behavior could be indicative of the solution not being fully converged at M1 = 0:10 because solution convergence is known to be dicult

to obtain at very low values of M1. In addition,

gures 21(a) to (n) allow for two other comparisons

between the two experimental data sets and the

two CFD solutions. The experimental data comparison in this low subsonic range (M1  0:24)

shows that the two sets roughly agree; this indicates a small Rn dependence but with the BART

Cp values generally having slightly more suction.

This result is consistent with 
ow at lower values

of Rn leading to earlier separation and reaching

higher suction peaks on wings with thin or sharp

leading edges. Because the CFD solutions were for

the test conditions, at least two of the test parameters are dierent between these sets: Rn diers

by a factor of 30 and M diers only slightly. The

result of the CFD comparison is that higher Cp suction peaks are predicted for 
ight than at BART

test conditions and these occur at or slightly outboard of the BART locations. This prediction is

not consistent with experimental experience, and

this reversal in the Rn trend is likely attributed to

the grid and turbulence model used. These modeling features worked well up to  = 10.4 but, as

already noted, do not yield the measured suction

peaks at  = 13.

Figure 21(o) presents Cp;u for the BART test as

determined from PSP (ref. 11) and corresponding

CFD. Even though there is some misalignment

distortion for the two halves of the gure, an

assessment of Cp;u shows that slightly more suction

is generated over the upper surface in the windtunnel test than predicted. Regardless of these

dierences, it is encouraging to note that some of

the key vortex-system features, such as location

of vortex origin and footprint, are captured by

both data types. (Note that even though the CFD

solution does not show the missile or its rail, they

were a part of the computation.)

Transonic. Four separate CFD transonic solutions were attempted (control surfaces unde
ected)

at the test parameters for FC 68, 69, 70, and 79

from table 5; however, only the one for FC 70

was run to convergence due to resource limitations.

Hence, there is only one transonic Cp comparison

presented for CFD and 
ight and it is contained in

gure 22.

Figure 22(r) shows the distribution of CFD and


ight Cp data at M1 = 0:97 and  = 4.4 (FC 70;


ight 152, run 5b); as before, the CFD solution

8
data serve as the background color mirrored about

the centerline. The 
ight port locations are de -

noted as black dots and the associated Cp values

by the color of the surrounding bubble outline. A

comparison of the data in gure 22(r), including

the insert, indicates very good overall agreement

by using this global Cp scale because the colors of

the bubble outlines are virtually indistinguishable

from the CFD over the wing. Figures 22(a) to (q)

show the detail agreement at both xed BL and

FS. As before, the detail comparisons are needed

because the global Cp scale at FC 70 is too large

for the entire upper surface of the airplane to capture the details that exist over just the wing alone.

Nevertheless, remarkable detail agreement occurs

all along the leading edge and along the chords

at BL of 40 and 55. Regarding the interest in

whether a shock crosses the hinge line of an aft

control surface could be detected in 
ight and predicted at BL = 153.5, a shock is noted to occur

near x=c = 0:75 in the 
ight data. (See g. 22(h).)

This location is aft of the aileron hinge line, and

this feature is not captured by the CFD of the con-

guration with unde
ected control surfaces.

The fact that the Cp agreement is better inboard and along the leading edge needs to be examined in light of the CFD assumptions that the


ight control surfaces were unde
ected and the

wing was rigid. A post
ight examination of the

angles of control surfaces for 
ight 152, run 5b

determined them to be for left elevon, 0.6; left

aileron, 1.8; right elevon, 2.0; right aileron, 1.8;

right leading-edge 
ap, 9.0; and left leadingedge 
ap, not available but assumed equal to that

on the right side; positive control-surface de
ection

is dened for any surface for which the 
ap edge is

down. Though these numbers for the trailing-edge

control-surface de
ections are  2, they can have

a powerful local eect on Cp and, consequently, on

wing shock location. Note the large disagreement

at the aftmost FS 450, 462.5, and 492.5 (gs. 22(o)

to (q)); this indicates a shock hitting the surface

near 2y=b = 0:4 and outboard of 2y=b = 0:5 for

FS 492.5. Regarding wing rigidity, during the entire 
ight test the outer wing was not as rigid as

the inboard wing, an expected occurrence; however, the eects at FC 70 should be minimal since

it is for 1g 
ight.

Contours. Figure 23 shows Cp upper surface

contours at M1 = 0:97 and  = 4.4 (FC 70,


ight 152, run 5b) for both 
ight and CFD. (CFD

contours are generated using FAST (ref. 23) and


ight contours from Tecplot (ref. 25).) There are

many similarities between these contours; overall,

the CFD pressures are more negative, especially

near the leading edge and over the aft part of the

wing. Along the inboard edges, aft about one third

of the distance from the S-blend curve, both the


ight and CFD data of gure 23 show a shock

followed by an expansion, whereas outboard only

the 
ight data show a shock aft centered near the

aileron hinge line. This elevon shock in 
ight has

been previously noted (g. 22(h)), and because it

does not show up in the CFD solution it is most

likely attributable to dierences in the controlsurface de
ections. The CFD modeling is with

the control surfaces unde
ected, whereas the 
ight

experiment has small, but measurable, trailingedge de
ections.

Boundary-Layer Proles at FC 7

Figures 24(a), (b), (c), and (d) show the comparison of measured and predicted boundary-layer

proles for rake locations 3, 4, 7, and 5, respectively, at FC 7. (See tables 5 and 12.) These locations were chosen because they had 
ows which

should be markedly dierent at nom = 13 and

subsonic speeds. At these respective positions,

the 
ows range from (a) being nearly streamwise,

(b) underneath the primary vortex, (c) at the secondary separation line, and (d) underneath the secondary vortex. Figure 24(a) also presents an estimate of the prole repeatability for rake 3 because

this position was used as a control, 
own with each

of the others, and had the most benign 
ow. As can

be seen, the experimental velocity ratios only have

a small deviation from one another, and the prole

is well estimated for y > 0:25. For y  0:25, the

measured prole develops more quickly near the

surface than predicted, even with the \wall function" option being used in CFL3D.

Underneath the primary vortex (rake 4 location), gure 24(b) shows qualitative agreement between the measured and predicted results but not

quantitative. In particular, for y  0:25 the predictions are less than measured, and for y > 0:25

the reverse is true. Moreover, both results indicate

a jet-type 
ow to commence at y > 0:3. Both


ows also show regions of quasi-linear variation

9
of velocity with y, indicative of being outside the

boundary layer and just into the in
uence of the

primary vortex.

In the vicinity of the secondary vortex there is

general disagreement between the measured and

predicted values, as shown in gures 24(c) and (d).

The predicted values are signicantly dierent,

whereas the measured ones look to be similar;

moreover, the measured values do not look like

what is seen at rake 3 or 4 in gures 24(a) or (b),

respectively. Regarding the EFD/CFD disagreement for rake 7, the predictions indicate that at

the originally estimated location of the secondary

separation line, the prole develops the edge velocity value only a small distance o the surface and

thereafter retains that level. This constancy is in

contrast to the measured values which only reach

edge velocity near the rake extreme. The measured

velocity is not asymptotic at the rake extreme; this

leads to the conclusion that the maximum velocity

has not been achieved at this location. The measured prole for rake 5, underneath the originally

estimated location of the secondary vortex, also

only achieves edge velocity near the rake extreme;

however, the predicted values are markedly dierent with jet-type 
ow velocities occurring near the

surface over most of the rake height. Comparing

only the EFD proles for rakes 7 and 5 in gure 25(a) shows the velocity distributions are very

similar. Although the plan was to use the results

of an initial CFD solution to measure two dierent boundary-layer proles associated with two different features of the secondary vortex over this

part of the wing, it is apparent that only one was

captured. The quasi-linear growth of velocity for

y > 0:5 for these proles is associated with vortices around these boundary-layer rakes because

the velocity eld produced outside a representative

vortex system core varies as 1=r.

Even though these predicted boundary-layer

proles were dierent than those measured, some

understanding of the local 
ow measured closer to

the surface may be gained by examining the nal

CFD solution. Figures 25(b) and (c) have been

prepared for that purpose. Figure 25(b) shows

the stagnation pressure contours at FS 295 and a

representation of these two boundary-layer rakes.

From this gure, one can see that the two rakes are

computationally located in between the secondary

vortex and a third vortex system, as indicated

by the streamwise ribbons, and are not at the

planned positions associated with the secondary

vortex. (Because the third vortex system is located

outboard of the secondary vortex, this system is

not called a tertiary vortex. Moreover a tertiary

vortex would be under more of the in
uence of the

secondary vortex than it would be of the primary

system.) Figure 25(c) locates the origin of the third

vortex system as coming from that portion of the


ow which crosses the apex in the S-blend curve

region and from there proceeds over the primary

vortex. Hence, this 
ow gets swept under the

primary and moves outboard where it remains in

the vicinity of the wing leading edge, inboard of the

crank. This 
ow interaction accounts for the third

vortex sense of rotation. Such a vortex system

is unexpected, not seen in experiments, and most

likely an artice of this CFD grid/solution.

Consider again the measured velocity proles of

rakes 7 and 5 in gure 25(a). Based on the close

proximity of these two rakes, only a single vortex

system outboard of the primary vortex can be con-

rmed and it is the secondary vortex. Moreover,

the actual secondary vortex must be more outboard, and most likely larger, than the predicted

one shown in gure 25(c). This 
ow feature is due

to both rakes being encompassed by the actual secondary vortex and the third vortex not really a

contributor.

Skin Friction at FC 7

Figure 26 provides the measured and predicted

cf values at FS 330 for similar conditions in 
ight

and for CFD. This gure can be used to locate

and assess the impact of the vortex systems because they produce high velocities on the surface

which are measured by the modied Preston tubes.

Qualitative agreement is shown because both data

sets have at least two regions of high cf , which

is indicative of primary and secondary vortices.

These vortices occur at BL values near one another, for example, 89 versus 84 for the primary

and 108 versus 103 for the secondary, with the

measured results given rst. The signicant differences for the primary vortex are (1) the more

restrictive range of BL over which the vortex in
uence occurs in 
ight, and (2) the greater levels of

cf reached; that is, measured levels are 39 percent

larger than predicted. For the secondary vortex,

the measured and predicted levels are comparable

10
but the predicted vortex covers a much narrower

range of BL. This result is consistent with the pre -

vious discussion of the predicted vortex systems

given in gures 25(b) and (c) because the presence

of the third vortex is expected to reduce the spanwise extent of the secondary vortex. Moreover, at

least two other cf plateaus or peaks are predicted

outboard of the secondary vortex, one at BL 107

and one very near the leading edge at BL 113.

The most inboard plateau or peak is more likely to

re
ect the presence of the predicted third vortex

than the one very near the leading edge because

at FS 295 the third vortex is located laterally midway between the secondary vortex and the leading

edge. (See g. 25(c).)

On and O Surface Flow Features at

FC 46

Figure 27 was constructed with Intelligent Light

FIELDVIEW software (ref. 26) and shows the fusion (overlaying) of surface tuft images from three


ight cameras with CFD surface streamlines and

vortex-core representation at M1 = 0:53 and

 = 10.4 (FC 46, 
ight 144, run 3b). In particular, gure 27(a) presents the combination of

the three 
ight-camera images projected onto a

grid representation of the airplane. Figure 27(b)

shows the CFD surface streamlines compare well

with these tuft images. Figure 27(c) presents isosurfaces of the stagnation pressure (PLOT3D) at

a value of 0.78 and represents the locations and

extent of the various airplane vortex systems. Figures 27(b) and (c), with transparency employed

for the vortex systems iso-surfaces, are combined

to form gure 27(d). As expected from the results

of the surface comparison, the vortex system is well

located with respect to the 
ight tufts.

EFD Basic Data and Comparisons

This section provides basic, low-speed windtunnel force and moment data for the F-16XL-1

airplane and provides comparisons of wind-tunnel

surface pressures, wind-tunnel on and o surface


ow features, and 
ight surface 
ow features.

Forces and Moment in Wind Tunnel

Figure 28 presents the longitudinal forces and

moment obtained on the 0.18-scaled model in the

Langley 30- by 60-Foot Tunnel. The data show

the model has well-behaved force characteristics

and is approximately neutrally stable over this

range of  and CL at low M1. Moreover, if the

aerodynamic center shift has been well controlled

by planform shaping, one could expect only slight

elevon de
ections to be needed for trimmed 
ight

from subsonic to supersonic speeds.

Surface Pressures in Wind Tunnel

The eects of model scale and Rn variation on

Cp data from dierent facilities at low speed are

presented along with transonic Cp contours in this

section.

Eect of model scale and Rn on Cp at low

speed. Figures 29 to 31 and gures 32 to 34 provide representative streamwise (constant BL) and

spanwise (constant FS), respectively, Cp data collected during the BART and 30- by 60-Foot Tunnel tests for the 0.04- and 0.18-scaled models at Rn

of 0:5 106, 1:0 106, 1:12 106, and 2:1 106

and  of 5, 13, and 20. This restricted set of

data was further limited to only those BL or FS

which had results (sometimes only on the upper

surface) at all four values of Rn. Tables 9 and 10

conrm that these models have comparable numbers of upper surface ports at the values of BL

and FS chosen. For reference, the resulting four

values of BL have three locations inboard and one

outboard of the air dam, and the two FS have locations well forward of and across the air dam. Two

expected eects occurred from all these results,

namely, little dierence between the two data sets

near Rn ' 1:0 106 and larger Rn eects on the

upper surface than the lower surface. A more detailed study of the remaining features begins with

the BART data (0:5 106  Rn  1:12 106).

For these three data sets, signicant dierences in

Cp occur on the upper surface. However, a general

assessment is that the suction peaks occur at the

same locations and the peak values increase with

Rn and .

A comparison of the Cp data from the BART

and 30- by 60-Foot Tunnel tests in gures 29 to 34

shows the general eect of increasing Rn from

1:0 106 to 2:1 106 is a higher suction peak

in Cp for  = 5 and 20 but lower peaks at

 = 13. A portion of this reversed trend at

 = 13 is attributable to the test results from

the 30- by 60-Foot Tunnel being at 12.5 rather

than the 13 in BART. Peak suction values of Cp

11
usually occur at the same location but the distri -

butions are markedly dierent. Dierences noted

between the two wind-tunnel data sets are not well

understood. Many factors, including dierences

in turbulence characteristics, 
ow quality, model

scale, and model installation between these two

tunnel tests would be expected to produce some of

the dierences noted. Therefore, even though Rn

is highest in the 30- by 60-Foot Tunnel, due to its

larger model, and the Cp peak values generally increase with Rn and , there can be signicant local

dierences pointing to model scale and Rn eects|

only a portion of which can be sorted. Beyond the

slight model asymmetry dierences mentioned previously for the 0.04-scaled model, data uncertainty

is an issue. The fact that the location of the peak

Cp remains essentially xed with changes in Rn

attests to the aerodynamic sharpness of the wing

leading edges of these models. (See ref. 27 for an

example of Cp peak location dependency on Rn

for a small-radius, rounded, leading edge.) The

more negative values of Cp with increasing Rn for

a xed  is attributed to the vortex system becoming more tightly organized or coherent which

leads to a stronger in
uence on the surface 
ow

and pressure eld.

Transonic Cp;u contours. Figure 35 is developed from the 0.11-scaled model data from 109 upper surface pressure ports (ref. 18), re
ected about

the centerline, from which the unscaled Cp;u contours were created by FAST. This gure shows

the eect of increasing  from 2:27 to 20:04

at M1 = 0:95 and is used to assess when the

shock crosses the trailing-edge hinge line. Between

 = 2:04 and 4.27, the coalescence of positive Cp

isobars is evidence of a shock crossing the aileron

hinge line. This shock strengthens on the aileron

hinge line and moves inboard onto the elevon surface by  = 6:54. Figure 35 shows that the

trailing-edge shock remains a prominent 
ow feature through  = 11:35, even in the presence

of the leading-edge vortex system. However, by

 = 15:44, the leading-edge vortex system and its

attendant cross-
ow features in a transonic 
ow

cause the trailing-edge shock to be pushed beyond

the last set of ports. With the distribution of pressure ports available, no strong shock forms across

the elevon hinge line over the entire range of .

The location of the shock near the hinge line of the

aileron is generally in qualitative agreement with

that determined from the transonic FC 70 
ight

results reported previously. (See g. 22(h).)

On and O Surfa ce Flow (Air Dams O)

in Wind Tunnel

Some examples of using FAST to fuse on and

o surface data are presented in gure 36 for

visual data taken in BART at low speeds. This

gure shows the superposition of vortex core and

reattachment point paths extracted from vaporscreen images2 onto surface oil-
ow images for

 = 10, 13, and 15 at Rn = 1 106 and

0:5 106 on the model with the air dams o. The

general trend is for the core projections to agree

well with the surface oil-
ow features and for the

reattachment projections to be slightly outboard

of the features. As expected, progressive inboard

movement of both projected paths occurs with

increasing , and only a small change in the surface

oil 
ows occurs with increasing Rn at a particular

 because of the small/sharp leading-edge radius.

Overall, the best agreement between the data sets

occurs for  = 15 and Rn = 0:5 106.

Improvements in the comparisons of these data

sets can only be achieved by two renements. The

rst is to improve the quality of the image data

collected by doing more thorough planning and

execution of the vapor-screen portion of the test.

The second is to improve the application of the

image-process algorithms.

Surface Flow in F light

Two examples of the fusion of image data are

given for upper surface 
ows on the F-16XL-1

airplane: tufts with liquid crystals (g. 37) and

tufts with oil 
ow (g. 38). All tuft and liquid

crystal image data were taken on the left wing with

the right wing reserved for oil 
ow. These images

were taken from two cameras mounted near the

top of the vertical tail, left and right side. Both

2 Some of the projec tion paths, as discussed in appendix A,

are not straight but disjointed. This is primarily associated with

the particular vapor-screen image used and, with four exceptions,

all images were of the instantaneous type. The exc eptions use d

the average image and were employed whe n the instantaneous

image did not contain all the vortex features. They occurred for

(1)  = 13 and Rn = 0:5 106 at FS 435; (2)  = 13 and

Rn = 1:0 106 at FS 225; (3)  = 15 and Rn = 1:0 106 at

FS 257.5; and (4)  = 15 and Rn = 1:0 106 at FS 350.

12
gures show that the tuft data are represented

by \vectors," solid-line segments with the same

orientation, origin, and length of the respective

trailing tufts. (See appendix B and ref. 28.)

Figure 37 shows the liquid crystals exhibit a

high shear (dark) region, which is caused by the

highly energetic vortex core and system scrubbing

the upper surface for the test condition of   13,

M1 = 0:28, and Rn = 47 106. Thus, the

liquid crystal data correspond well with the 
ow

features detailed by the tuft data. (See section

\Skin Friction" for a discussion of corresponding

on and o surface 
ow.)

Figure 38 displays tuft data mirrored about the

aircraft centerline, \vectorized," and then overlayed onto a portion of the right wing with the

oil 
ow, also at   13, M1 = 0:28, and

Rn = 47 106. Good agreement is observed between these two types of surface 
ow techniques in

the highlighted and enlarged region shown at the

right.

Concluding Remarks

This report has focused on the comparative efforts regarding data obtained from 
ight, computational 
uid dynamics (CFD), and wind-tunnel

studies for the F-16XL-1 airplane supported, wherever possible, with the aid of data fusion (overlaying). The resulting highly diverse set of data types

was obtained over a wide range of test conditions.

From these comparative studies and the basic data

taken, the following remarks and observations can

be made:

1. The aircraft geometry compared closely with

the numerical surface description, diering by

at most 0.24 in. over the wing; one exception

was over the airplane forebody where the differences were larger and reached a maximum

value of 1.65 in. near the nose. However, even

this dierence leads to a nose droop angle of

only 1, which should have minimal impact on

the pressure results.

2. Overall upper surface pressure coecient (Cp;u)

comparisons of CFD and 
ight results have

been expedited by use of computational graphical tools.

3. At subsonic speeds, the 
ight suction peaks and

distributions were generally well predicted at

 = 5:5 and 10.4 (where  is angle of attack),

but underpredicted at nom = 13 and 20

(where nom is nominal angle of attack).

4. At transonic speeds, generally good pressure

agreement was noted for 
ight and CFD at

 = 4:4 especially near the leading edge inboard of the crank. The eect of small trailingedge control-surface de
ections ( 2) in 
ight,

not modeled with CFD, may have led to poorer

agreement for the aft pressures and to an unpredicted shock.

5. The location of the shock near the hinge line

of the aileron, as determined from wind-tunnel

Cp;u contours near  = 4 at a free-stream

Mach number (M1) of 0.95, is generally in

qualitative agreement with that determined

from the transonic 
ight results at  = 4:4 and

M1 = 0:97.

6. Turbulent signal output from hot-lm gauges

demonstrated that the 
ight leading-edge attachment line was turbulent; this makes this

data set in agreement with other 
ight data and

does not support the anticipated relaminarization of this boundary layer under the in
uence

of high suction peaks near the leading edge.

7. The grid used in the CFL3D code produced

Reynolds number like term for 
at-plate turbulent boundary layer (y+) values much larger

than 2 at 
ight conditions; in fact they were

 82. The \wall function" option in the code

proved sucient at subsonic speeds to represent the underneath attached 
ow and the primary vortex of the boundary layers but insuf-

cient underneath the secondary vortex system

at nom = 13. Moreover, the 
ight data indicate the presence of an interaction of the secondary vortex system and boundary layer.

8. Qualitative agreement is indicated between predicted and measured skin friction spanwise distributions at subsonic speeds because both CFD

and 
ight results show at least two regions of

high local skin friction coecient (cf), which

is indicative of primary and secondary vortices

at spanwise positions near one another. However, the absolute levels of cf and the indicated

spanwise extent of the vortices dier.

9. Data fusion (overlaying) has been applied to

a variety of 
ow-visualization techniques for

13

ight, wind tunnel, and CFD with insightful re -

sults. In particular, combined surface tuft images from three 
ight cameras compared with

CFD surface streamlines and vortex-core representation show generally good agreement at

M1 = 0:53 and  = 10:4, and the 
ight results

of oil 
ow with tufts and liquid crystals with tufts

compared well, but the wind-tunnel comparison of

vortex core and reattachment locations|as determined from vapor screen and point compared with

surface oil-
ow traces|though showing promise,

point to a need for further renement.

14
Appendix A

Visualization Tools

Four dierent visualization tools are discussed|

ILLUME, MAPPER, Core- and ReattachmentPoint Locator, and Tuft Representer. An explanation of each tool and its function are as follows.

ILLUME

Light-sheet 
ow visualization experiments performed in wind tunnels or in 
ight require con-

guring a recording camera to capture the experiment and a light sheet to illuminate particles in

the 
ow eld. Often valuable experimental time is

expended trying to determine the optimal placement of the camera and light sheet. The software

package named ILLUME (Interactive Light Sheet

Locator Utility and Modeling Environment) allows

for the computer simulation of either experimental

wind-tunnel or 
ight test camera and light sheet

locations, in combination with the geometry of interest, in order to ensure a high degree of visualization success before the start of the test. ILLUME

is interactive and allows the user to simulate the

position and orientation of one or more cameras

and one or more light sheets while viewing, from a

selected camera, the intersection of these elements.

For each camera, the user can simulate the eect of

changing the lens on the eld of view to determine

the preferred one from those available for the test.

ILLUME can also be used, as done for this test,

to view the geometry and any surface features of interest. This viewing could include those representing the location of the video targets. A sucient

number ( 6) of viewable video targets per video

camera are required in order to establish the camera position and orientation; this process is called

space resection in reference 29. This can lead to the

establishment of 
ow features as quantiable data

when the video images are projected to the surface. Additional information about ILLUME can

be found in reference 30.

MAPPER

MAPPER is an updated version of a software package called Reconstructor (refs. 27 and

31), which enables a 2-D visual image, existing

at a camera image plane, to be projected onto

the 3-D surface where it occurred. This process has been applied to both wind-tunnel vaporscreen and 
ight surface-
ow (tuft) images in

these references. (These papers are retrievable

from the NASA Langley Technical Report Server

at http:==techreports:larc:nasa:gov=ltrs=ltrs:html:)

Additional information about the use of this process is also available from reference 32.

Core- and Reattachment-Point Locator

The Core- and Reattachment-Point Locator

software tool is an update of the Extractor package

described in reference 27. Its use is to locate the

vortex-core and approximate reattachment-point

positions from vapor-screen images obtained in the

BART. It can be exercised from UNIX platforms

using the Precision Visuals Workstation Analysis

and Visualization Environment (PV WAVE) software system.

Tuft Repre senter

The Tuft-Representer software tool depicts individual tufts in a tuft image as a straight-line segment. The origin is on the surface at the base of the

free trailing tuft with the segment pointing along

the average direction of the trailing tuft. With the

characteristics of origin and direction, this representation is a vector (without arrowhead) in nature

and is so referred to in that manner. This tool operates on a Sun workstation and requires the locations of the video targets and camera parameters

as input data. (Flight values of video targets and a

schematic representation are given in table A1 and

g. 7, respectively.) With the use of MAPPER,

this vector representation can be projected onto

the upper surface of the airplane. Additional details about the Tuft Representer can be found in

reference 20.

15
Table A1. Video Targets Measured at Dryden Flight Research Center on F-16XL-1 Airplane

Target FS BL WL

1 159.712 24.391 102.008

2 163.961 38.667 90.697

3 174.118 30.260 95.106

4 172.851 40.776 91.630

5 187.337 20.662 111.121

6 182.924 39.493 93.194

7 182.040 53.637 90.917

8 195.672 41.192 94.583

9 195.025 52.305 92.941

10 202.439 11.686 120.854

11 218.023 21.643 109.554

12 209.345 54.023 94.248

13 207.507 66.298 91.365

14 255.566 9.497 118.915

15 276.779 20.845 108.481

16 275.204 37.794 99.424

17 254.526 71.183 94.841

18 268.979 90.684 89.992

19 282.760 80.433 94.617

20 319.019 70.367 96.999

21 311.541 94.559 93.405

22 305.890 103.870 89.276

23 354.705 35.863 100.601

24 345.123 71.892 96.859

25 327.117 111.547 88.813

26 353.944 90.010 95.659

27 345.800 118.396 88.436

28 383.740 62.037 96.447

29 380.734 95.101 95.466

30 378.884 118.190 92.682

31 402.714 38.917 101.689

32 403.856 72.841 95.834

33 401.002 103.350 94.970

34 398.664 126.490 92.244

35 431.257 69.354 95.169

36 429.992 93.525 94.958

37 433.816 115.440 94.295

38 452.936 67.818 94.603

39 449.874 84.811 94.600

40 450.587 101.896 94.438

41 466.305 82.975 94.283

42 460.331 124.085 97.253

43 470.263 179.820 93.004

44 457.041 197.147 91.778

45 440.980 197.296 90.903

46 423.935 198.037 90.136

47 202.557 26.507 73.928

48 249.466 28.655 76.361

49 276.735 28.553 79.282

16
Appendix B

Data Processes and Management

Two main topics are covered in this appendix: the various processes used to get the data from its

as-received or reduced form to that needed for display and the management of the data and its meta-data.

These two topics have been handled in an integrated manner from near the beginning of this project.

Data Processes

Data from experimental tests in wind tunnels and 
ight along with predicted values from CFD all

require some processing in order to be utilized in a comprehensive way. They are discussed in order.

After dening a common data format|TOAD (ref. 33)|for the pressure and force and moment data

and employing data translators to bring all experimental results into that format, the les of pressure data

generally required two other processes applied before they could be used. The rst was, if necessary, to split

the pressure les of complete upper and lower surface data into separate les, split flight or split BART,

and the second was to slightly alter the coordinates of the ports from actual to nominal values, act to nom,

in order that the Cp values be directly comparable with other data sets. These two processes, along with

the TOAD le plotting package, tplot, can be exercised from UNIX platforms by using the Precision

Visuals Workstation Analysis and Visualization Environment (PV WAVE) software system.

Additional antecedent processes were required in order to get the 
ight pressure time histories, collected

at the Dryden Flight Research Center, ready for use in the preceding processes. First, a particular time

interval from the 
ight time history had to be established by examining the video data from the three

cameras to determine a 2-sec segment when acceptable image data were recorded from all cameras. Second,

the pressure les were processed at Dryden for these times. Third, these les were transferred to Langley

where they were rewritten into the TOAD format. Also transferred were selected parameter statistics and

the associated pressure le name so that a summary report (meta-le) could be created. Fourth, these

summary reports were processed locally on a Silicon Graphics workstation using the nonnumeric TOAD

editor procedure, nnted, a variation of the TOAD editor procedure (ref. 34), ted, that can accommodate

nonnumeric data in printing, and the results sent to the data management system for automatic data-eld

population. Only the items related to Test Technique were added to the database for each 
ight/run.

Two processes were required to generate the needed CFD Cp les. First, the software package

FIELDVIEW (ref. 26) was utilized to cut the entire solution at the surface for selected BL and FS values

on both the upper and lower surfaces and to save the resulting les. Second, these les were translated

into TOAD format for subsequent use.

Data Management

System. The Program for Information Storage and Management (PrISM) (ref. 35) is a graphical,

direct-manipulation Structured Query Language (SQL) interpreter. Users are able to query, add to, and

modify data in a database without needing to know the internal representation of the data as tables and

columns.

PrISM consists of a graphical user interface, database interface functions, and PrISM internal database

tables. The PrISM internal database tables consist of core tables, which are constant and unchanging

across all PrISM databases; kernel tables, the values of which depend upon the specic PrISM database

and determine the PrISM functionality available to users; and tables which support the PrISM Security

System.

The application of PrISM to the F-16XL-1 High-Lift Project and the Cranked-Arrow Wing Aerodynamics Project (CAWAP) involves the storage of information about 
ight tests, wind-tunnel tests, and

CFD solutions. The database stores administrative information about the tests (meta-data) and the full,

mass storage, path name of the resulting data les. For EFD measurements, the experimental data stored

17
are les of pressure, force/moment, still photographs, and two-dimensional images digitized from videotape

(both instantaneous and time averaged); for CFD predictions, the grid and solution les are stored as well

as Cp data at selected FSs and BLs. The F-16XL-1 database also supports the viewing of 3-D renderings of the 2-D 
ight images through MAPPER. Queries for selected pressure data and 2-D 
ight image

data are available to WWW users by completing the appropriate form under Program for Information

Storage and Management, F16XL High-Lift Project CAWAP given in reference 36. As a result of

performing such a query, data retrievals are possible. For example, depending on the selection made, one

may have either (1) the contents of selected pressure les viewed through hypertext links and/or the les

retrieved and made local for further processing or (2) the 2-D 
ight image les viewed through hypertext

links.

Pressure le naming convention. Designations of particular F-16XL components making up a

conguration tested, the facility used in its testing, the research engineer, and/or the test conditions

may appear in the names of the pressure les saved on mass storage. Three basic F-16XL wind-tunnel

congurations have been tested and they are designated as b1, b2, and b3 congurations. Conguration b1

is the model with air dams, b2 is the model with air dams and missiles, and b3 is the model with missiles.

At Langley, low-speed tests were made for all three congurations in BART but only for congurations b2

and b3 in the Langley 30- by 60-Foot Tunnel. At the Ames Research Center, high-speed tests were made

for conguration b2 in the Ames 11-Foot Tunnel.

The le names used in the keys in the presented pressure data plots dier from the names by

which they have been saved on mass storage in four ways: (1) they have been simplied by removing nonvital information, (2) they have been simplied by removing test or solution information,

which has been placed in the gure caption, (3) they re
ect the wind-tunnel facility and not the

test engineer, and (4) they are pressure surface specic as a result of processing the saved le. Examples of the changes to the le names for wind-tunnel and 
ight are as follows: the le name

\dmss:@jelama/BART/press.bart.Q40.b2.basic" is split into two les \BART,up" and \BART,low"

for the basic F-16XL model with air dams and missiles (conguration b2) at q1 = 40 lb/ft2

or M1 = 0:17 and Rn = 1:12 106; the le \dmss:@jelama/hahne/press.409.b2.basic" becomes

\30 by 60,up" for the same conguration at q1 = 6:7 lb/ft2 or M1  0:07; and the 
ight le

\dmss:@jelama/f16XL.
ight/
ight144/
t0144.tp16b.cp.toad" is split into two les \Flt144.Run16b,up"

and \Flt144.Run16b,low," thereby, uniquely identifying the 
ight-and-run and consequently the specic

nominal 
ight conditions from table 5 for the same conguration. For certain wind-tunnel data comparisons

with an Rn variation, the simplied lename will have a sux added, e.g., \R2.1" to signify Rn = 2:1106.

A CFD example is le \dmss:@jelama/garriz/F16XL/aoa19.84m0.242/BL55.up.toad" being reported as

\CFL3D" at FS 55 and the actual 
ight conditions of  = 19:8 and M1 = 0:24.

Files of 
ight, nominalized upper and lower surface pressures have also been saved on mass storage and

have a naming convention typied by \dmss:@jelama/f16XL.
ight/
ight144/
t0144.tp16b.cp.up tmp" and

\dmss:@jelama/f16XL.
ight/
ight144/
t0144.tp16b.cp.low tmp," respectively.

18
Appendix C

Airplane Instrumentation

Due to the complexity of the airborne research

systems for the basic airplane, the 
ight testing

was divided into three parts, designated \Parts 1A

to 1C." These were ordered according to increased

levels of instrumentation complexity. Part 1A

was the initial phase and had a limited but su-

cient instrumentation suite for airplane documentation at the Langley Research Center. The suite

also included video cameras to capture the images from the HUD and some surface-
ow features. Part 1B, also performed at Langley, was

done with additional instrumentation for both airplane positioning and surface-
ow studies and included additional video cameras. Part 1C, later

known as CAWAP, was performed at the Dryden

Flight Research Center and had the most extensive instrumentation suite for detail surface-
ow

measurements. This suite included instrumentation to measure the pressures from 
ush and

belted static ports, boundary-layer rakes, modi-

ed Preston tubes, and the electrical signals from

a hot-lm belt. (See g. 6.) Details of the instrumentation suites, the modications made, and


ight practices with them are given in the following

sections.

Instrumentation Suite

The core of the airborne instrumentation systems was the Airborne Test Instrumentation System (ATIS). The ATIS had associated with it either a rst- or second-generation System Control

Unit (SCU) so that the nal designation is either ATIS-SCU1 or ATIS-SCU2. The ATIS-SCU1

is a 10-bit (5.12 to +5.12 V) pulse-code modulated (PCM) modular data system and was installed originally by General Dynamics for USAF


ight testing. For the Cranked-Arrow Wing Aerodynamics Pro ject (CAWAP), ATIS-SCU2, a 12-bit

(10.24 to +10.24 V) PCM was used and the system reprogrammed to accommodate the measurements list. Some airplane parameters proved to

be from dicult to impossible to recalibrate after

the upgrade from the 10-bit to the 12-bit system;

therefore, the existing 10-bit calibrations were used

but only after rescaling the parameter voltages

(and engineering units) to 12-bit decimal counts.

For example, if a 10-bit system channel was programmed with Oset = 0 and Gain = 1, 5.12 V

would correspond to 0 counts and +5.12 V to 1023

counts. In the 12-bit system, 10.24 V corresponds to 0 counts and +10.24 V to 4095 counts.

Therefore, a parameter whose voltage output for a

given engineering unit value is 2.56 V would equal

256 counts in a 10-bit system but 1536 counts in

a 12-bit system. Zero oset and gain settings for

each PCM channel were taken into consideration

when doing calibration conversions. The parameters that were corrected in this manner included

pitch, roll, and yaw as well as rudder, elevon, and

aileron forces and hinge moments (which all together accounted for 16 measured parameters). All

other airplane parameters were recalibrated using

the 12-bit data system. The ATIS also provided

time code for insertion into the ve airborne video

cassette recorders for synchronization of data in

the CAWAP.

Mod ications

The specic research modications relative

to the basic airplane instrumentation suite for

the preliminary on surface-
ow measurements,

Part 1A, were as follows:

External paint scheme for on and o surface


ow visualization

Externally mounted lipstick cameras on either

side at the top of the vertical tail with associated VCRs and time code inserters, one for

every two VCRs, in the forward equipment bay

Video camera for the HUD and associated VCR

and time code inserter in the forward equipment bay

Enclosure of the lower surface AMRAAM missile cavities

Enclosure of the gun muzzle trough

The external paint scheme (g. 1(a)) was based

on the scheme used on the Dryden F-18 HighAngle-of-Attack Research Vehicle. The upper

surfaces of the airplane were painted 
at black to

provide sucient contrast with the tufts, oil, and

liquid crystals so that the on surface 
ow patterns

could be visualized in Parts 1A and 1B. The gold

stripes on the wing were used to identify the leading edge in the video images and were to be of particular value for the o surface 
ow-visualization

portion of the test, which did not occur.

Two tail-mounted video cameras (left tail camera shown in g. C1) were used to provide coverage

over the upper surfaces of the wings. The locations of these cameras are shown in gure 12. The

elds of view of these cameras could be changed

by three methods: (1) installing lenses of dierent focal length, (2) adjusting the camera pivot

angle (angle of the camera centerline with respect

19
to vertical) up to 25, and (3) installing one of

four camera mounting blocks, which provided cant

angles (angle between the camera centerline and

the longitudinal axis of the airplane) of 8.5, 17.5,

26, and 36. Lenses used in these locations were

almost exclusively 7.5 mm.

The AMRAAM lower surface cavities were covered with plates to produce a 
at surface. The gun

muzzle trough was smoothed over with a plug to

preclude the formation of extraneous vortices not

associated with the basic cranked-arrow wing planform. The gun trough is shown with and without

the plug in gure C2.

Part 1B began with the instrumentation suite

of Part 1A and added other items to bring it to the

required level. The changes included

Modications to the ammo drum bay instrumentation pallet

Global Positioning System (GPS)

Wingtip video cameras installed in dummy missiles noses (2)

Fuselage-mounted video cameras (2)

Enhanced measurements of airplane state

parameters (linear accelerometers at the airplane center of gravity, attitude gyros, rate

gyros, 
ow vanes, and nose-boom pitot-static

transducers)

Engine instrumentation

Tail-mounted microphones (3)

An instrumentation pallet, based on a design

provided by the Air Force Armament Laboratory,

was installed in the ammo drum bay which is an

unpressurized compartment in the fuselage located

aft of the cockpit. The pallet, shown in gure C3,

could be raised while installed in the airplane to

permit access to components located on the lower

shelves. The pallet contained components associated with the GPS, the measurement of the enhanced airplane parameters, and the microphones.

It also contained the two additional VCRs added

for Part 1B and their associated time code inserter.

The GPS was installed to provide accurate realtime airplane position and to improve the accuracy

of the airspeed calibrations. Because of the limited

internal volumes in the airplane, the components

were distributed throughout the vehicle including

on the ammo drum bay pallet.

For Part 1B, there were seven video cameras

and ve VCRs, an increase of four and two, respectively, over Part 1A. These were added to provide a more complete video coverage of the wing

apex regions and leading edges. Note that the apex

regions were viewed by lipstick cameras mounted

externally on the fuselage (g. C4), just behind

and to the side of the cockpit, whereas the cameras used to view the leading edges were mounted

in the heads (g. C5) of dummy, but properly mass

distributed and ballasted, AIM-9L missile bodies.

The fuselage cameras had 4-mm lenses and were

congured such that they could feed video signals

to either the VCRs located in the forward equipment bay or the ammo bay, depending on the cameras needed for a particular 
ight. The missile

cameras had 15-mm lenses and were congured to

only feed video signals to the VCRs in the ammo

bay. (Table C1 summarizes the nominal center of

lens camera locations for the video cameras, and

table C2 summarizes both lens location and orientation, as determined by resection (ref. 29), for the

CAWAP 
ights. The camera angles are dened in

g. C6.)

The CAWAP instrumentation suite began with

that of Part 1B and other items were added to

bring it to the required level. The changes occurred

because of the need to measure the following items:

Pressures for 
ush and belted static ports

Pressures from boundary-layer rakes

Pressures from modied Preston tubes

Electrical signals from a hot-lm belt

Practices

Pressure location measurements and assignments. In order to compare the pressure measurements from the 
ight experiments with CFD

predictions and wind-tunnel model measurements,

the ports on the airplane were distributed in the

same relative locations as those on the wind-tunnel

models. Becasue the airplane had more available

pressure measurements, additional ports were included in both spanwise and chordwise distributions. In particular, 
ush ports were located in the

leading-edge region of the right main wing with the

pressure belts being used to complement these locations on the main and outboard right wing panels. The pressure belts, composed of multiples of

10 tube subbelts, were necessary because the wing

contains the fuel and cannot be pierced except at a

few cavity locations. Care was taken in the selection and installation of the belts in order to minimize the eect of the belts on the measured pressures. Tubing diameter was kept at a minimum by

maintaining short belt lengths to the appropriate

pressure transducer. In addition, noninstrumented

20
tubing was placed outside of the tubes with ports,

and then ramps from the wing surface to the top of

the tubes were faired in. Figure C7 shows a crosssectional sketch of the tubing and ramp layout on

the wing. Similarly, a single ramp was added to

the forward end of the tubes, in the vicinity of the

leading edge, in order that the 
ow be disturbed

as little as possible, as shown in the photograph of

gure C7.

The rst step in placing the ports on the wing

was to identify the desired locations by measuring

from known reference points on the wing. After the

pressure belts were installed, the static-pressure

measurement (port) holes were punched into the

belt material at the desired locations. Some ports

had to be relocated because of wing surface conditions. Once the ports were in place, a second measurement was made by using a theodolite method

to determine the exact location of each port on the

wing surface. Since the geometry of the airplane

was rst measured with photogrammetry, leadingedge and trailing-edge locations and chord lengths

were determined from these measurements.

Dierences were noted between the specied

port locations and those laid out. The dierences

were due to keeping the pressure belts streamwise,

which makes the BL values vary along the FS, and

the physical constraints associated with the underlying structure in the vicinity of the wing apex and

leading edge. Therefore, these measurements had

to be adjusted to nominal values after installation

in order that comparison plotting could be done.

For the belts it was just a matter of changing the

BL value from actual to nominal. However, for

the 
ush ports along the leading edge and in the

apex region, an adjustment in the actual BL and

FS was needed. In order to make these 
ush port

adjustments, each port was assigned a nominal FS

and BL based on the actual value of the FS and

the desired BL. The FS for the 
ush ports was established based on the intersection of the desired

BL with the wing leading edge. Table 4 lists the

actual locations as measured and the new nominal

locations.

Video. Prior to each 
ight, the researchers selected the ve video cameras whose signals were

to be recorded. Typically, the HUD camera was

recorded only during airspeed calibrations and the

aeroacoustics and handling qualities 
ights. During the 
ow-visualization 
ights, the outputs from

as many as ve external cameras were recorded.

(See g. 12.)

For CAWAP, an initial video recording was

made on the right wing to establish the status of

the pressure belts. The cameras were located on

the right side of the airplane atop the vertical tail

and on the fuselage. During later 
ights, the focus

was the wing surface 
ow eld as depicted by the

tuft patterns on the left wing. These patterns were

video recorded from the cameras located atop the

left vertical tail, on the fuselage, and in the left

missile nose.

21
Table C1. Nominal Center of Lens Camera Locations for Part 1B an d CAWAP

Camera position FS BL WL Lens, mm

Left tail 563.322 2.042 223.058 7.5

Left fuselage 181.24 15.9 119.7 4

Left missile 403.68 201.046 91.74 15

Right tail 563.322 2.042 223.058 7.5

Right fuselage 181.24 15.9 119.7 4

Right missile 403.68 201.046 91.74 15

Table C2. CAWAP Center of Lens Camera Locations and Orientations as Determined by Resection

Camera position FS BL WL !, deg , deg , deg Lens, mm

Left tail 559.309 2.044 226.817 131.800 44.528 121.073 7.8223

Left fuselage 181.388 19.125 117.798 149.290 3.384 154.513 3.9106

Left missile 408.016 206.524 89.185 1.082 44.565 0.795 15.877

22
Camera

Close-up view

Figure C1. Video camera installation on vertical tail (left side) of F-16XL-1 airplane.

23
(a) Without plug.

(b) With plug.

Figure C2. Gun muzzle with and without plug.

24
Ammo drum bay pallet

Figure C3. Ammo drum bay instrumentation pallet.

Camera

Figure C4. Fuselage -mounted video camera used for viewing leading-edge apex region of wing.

25
Camera

Figure C5. Dummy AIM-9L Sidewinder missile used as wingtip video camera pod.

26
κ

ω φ

x y

z

Camera

Figure C6. Euler angles used for airplane cameras.

27
50:1 Fairing Dummy tubes Active tubes

3.1 0.0626 O.D.

Figure C7. Sketch of pressure belt cross section showing active tubes and ramps. Dimensions are in inches.

28
Appendix D

Description of Wind-Tunnel Test and

Vapor-Screen Systems

In support of this project, three Langley windtunnel tests of F-16XL models were conducted, one

in the Langley 30- by 60-Foot Tunnel (ref. 37) as

test 50 in 1994 and two in the Langley Basic Aerodynamics Research Tunnel (ref. 38) as tests 3 and

4 also in 1994. A brief description of tests 50 and

3 follow. Test 4 is described in reference 11 and it

was a pressure sensitive paint study of the upper

surface with no air dams, which has no counterpart

in 
ight or CFD.

Test in Langley 30- by 60-Foot Tunnel

This open-jet wind-tunnel force/moment and

surface pressure measurement test 50 was conducted in the 30- by 60-Foot Tunnel with a 0.18-

scaled model mounted on a six-component balance

in an aft-sting arrangement (g. D1) at M1  0:07

or q1 = 6:7 lb/ft2 and Rn = 2:1 106. The model

was tested at  = 5:0, 2:5, 0.0, 2.5, 5.0,

7.5, 10.0, 12.5, 15.0, 17.5, 20.0, 22.5, 25.0,

27.5, and 30.0 and at  = 20:0, 8:0, 2:5,

0.0, 2.5, 8.0, and 20.0. Various combinations

of control-surface de
ections and other geometrical

variations of the baseline model were tested. Only

the results for the baseline model at  = 0 are reported because no other 
ight or CFD results are

available for comparison. (The pressure port locations on this model are nominally represented on

the airplane by a subset. Compare tables 9 and 4.)

For reference, o surface 
ow results were obtained by using the vapor-screen technique. The


ow was seeded with propylene diglycol vapor,

illuminated by a nearly normal laser sheet (located far outside the tunnel 
ow) sweeping over

the right half of the model in a small arc, and

recorded with a xed video camera located behind

the model. However, because of the diculty in resolving the camera parameters, these results could

only be considered qualitative and therefore are not

presented.

Test in BART

BART has an open return and a test section made of Rohm and Haas Plexiglas material. The test 3 was conducted with a 0.04-scaled

model, mounted on a single pedestal (g. D2), for

ranges of M1, q1, and Rn from 0:07 to 0:17,

8:4 lb/ft2 to 40 lb/ft2, and 0:50106 to 1:12106,

respectively. The model was tested at corrected

values of , which took into account the in
uence

of the sidewall pressures determined at start of test.

The test values of  were 0.0, 5.0, 10.0, 13.0,

15.0, and 20.0 at  = 0. No control surfaces

were de
ected but dierent model components and

geometrical variations were examined. In addition

to surface pressures, an oil-
ow study was done on

the surface for the same values of . The pressure port locations on this model are nominally

represented on the airplane by a subset. (Compare tables 10 and 4.) O surface studies using

a ve-hole probe and vapor screen were done over

a restricted set of the test ; namely, 10, 13,

and 15. This was done to emphasize the takeo and landing portion for the High Speed Civil

Transport study. The ve-hole probe data were

collected within survey boxes which used the model

right wing upper numerical surface description as

a lower boundary at two fuselage stations, which

have full-scale values of FS 225 and FS 350. The

other boundaries for the two boxes were BL 41.5

and BL 50.625 inboard, BL 92.15 and BL 145.55

outboard, WL 125 and WL 150 top, and WL 87.5

and WL 87.5 bottom at FS 225 and FS 350, respectively. The vapor-screen studies were done by

seeding the 
ow with propylene digylcol vapor, illuminating the 
ow with a translating laser sheet

(mounted outside and atop the tunnel) shining on

the top of the model, and recording images with a

xed video camera mounted atop and ahead of the

test section. Images were taken at seven fuselage

stations which have full-scale values of FS 167.5,

FS 225, FS 257.5, FS 300, FS 350, FS 390, and

FS 435. Sample images are shown in gure D3.

29
Figure D1. 0.18-scaled model of F-16XL in Langley 30- by 60-Foot Wind Tunnel.

30
Figure D2. 0.04-scaled model of F-16XL in Langley BART.

31
Left wing

primary

vottex

FS 225

FS 300

FS 435

Figure D3. Sample vapor-screen images for 0.04-scaled model of F-16XL airplane in Langley BART at

 = 15, M1 = 0:15, and Rn = 1 106.

32
References

1. U.S. Standard Atmosphere, 1962. NASA, U.S. Air Force, and U.S. Weather Bureau, Dec. 1962.

2. Bertin, John J.; and Smith, Michael L.: Aerodynamics for Engineers, Second ed. Prentice-Hall, Inc.,

1989, p. 146.

3. Ghaari, Farhad: Navier-Stokes, Flight, and Wind Tunnel Flow Analysis for the F/A-18 Aircraft.

NASA TP-3478, 1994.

4. Potsdam, M. A.; Intemann, G. A.; Frink, N. T.; Campbell, R. L.; Smith, L. A.; and Pirzadeh, S.:

Wing/Pylon Fillet Design Using Unstructured Mesh Euler Solvers. AIAA-93-3500, Aug. 1993.

5. Hillaker, H. J.: F-16XL Flight Test Program Overview. AIAA-83-2730, Nov. 1983.

6. Talty, Patrick K.; and Caughlin, Donald J.: F-16XL Demonstrates New Capabilities in Flight Test at

Edwards Air Force Base. J. Aircr., vol. 25, no. 3, Mar. 1988, pp. 206{215.

7. Anders, Scott G.; and Fischer, Michael C.: F-16XL-2 Supersonic Laminar Flow Control Flight Test

Experiment. NASA/TP-1999-209683, 1999.

8. Bower, J. N.; and Scott, S. R.: The F-16XL Flight Test Program. Society of Flight Test Engineers

15th Annual Symposium, Aug. 1984, pp. 9-1{9-5.

9. Bates, J. D.; and Livengood, J. T.: F-16E Basic Lines Data Report|Volume 1. Rep. no. 400PR011,

General Dynamics Corp., Dec. 1981.

10. Bates, J. D.; and Livengood, J. T.: F-16E Basic Lines Data Report|Volume 2. Rep. no. 400PR011,

General Dynamics Corp., Jan. 1982.

11. Rickard, Susan J.; Washburn, Anthony E.; Morris, Martin J.; and Donovan, John F.: Pressure Sensitive

Paint Studies of Vortical Flow at Low Speed. SAE Paper 951989, Sept. 1995.

12. Bertelrud, A.: Total Head/Static Measurements of Skin Friction and Surface Pressure. AIAA J.,

vol. 15, no. 3, Mar. 1977, pp. 436{438.

13. Thomas, J. L.; Krist, S. T.; and Anderson, K. W.: Navier-Stokes Computations of Vortical Flows Over

Low-Aspect-Ratio Wings. AIAA J., vol. 28, no. 2, Feb. 1990, pp. 205{212.

14. Thomas, James L.; Weston, Robert P.; Luckring, James M.; Walters, Robert W.; Reu, Taekyu; and

Ghaari, Farhad: A Patched-Grid Algorithm for Complex Congurations Directed Towards the F-18

Aircraft. AIAA-89-0121, 1989.

15. Goldstein, Richard J., ed.: Fluid Mechanics Measurements. Hemisphere Publ. Corp., 1983,

pp. 565{569.

16. Fisher, David F.; and Fischer, Michael C.: Development Flight Tests of Jetstar LFC Leading-Edge

Flight Test Experiment. Research in Natural Laminar Flow and Laminar-Flow Control, Jerry N.

Hefner and Frances E. Sabo, compilers, NASA CP-2487, Pt. 1, 1987, pp. 117{140.

17. Spellman, M. W.: Model and Test Information Report 1/9-Scale F-16E Force and Loads Model. Rep.

no. 400PR026, General Dynamics Corp., Aug. 1991.

18. Elbers, W. K.: Wind Tunnel Data Report 1/9-Scale F-16E Pressure Model NASA Ames Research

Center Tests 517-1-11 and 517-1-97. Rep. no. 400PR037, Vol. II, General Dynamics Corp., Dec. 1981.

19. Hahne, David E.: Low-Speed Aerodynamic Data for an 0.18-Scale Model of an F-16XL With Various

Leading-Edge Modications. NASA/TM-1999-209703, 1999.

20. Lessard, Wendy B.: Subsonic Analysis of 0.04-Scale F-16XL Models Using an Unstructured Euler

Code. NASA TP-3597, 1996.

21. White, Frank M.: Viscous Fluid Flow. McGraw-Hill, Inc., 1974, pp. 474{476.

33
22. IBM Visualization Data Explorer Version 3.1 Users Guide, Sixth ed. Thomas J. Watson Res. Center

Publ. No. SC38-0496-05, IBM Corp., 1995.

23. Walatka, Pamela P.; Clucas, Jean; McCabe, R. Kevin; Potter, Rick; and FAST programmers at ARC:

FAST User Guide: Version 1.1a. RND-3-010, NASA Ames Research Center, Aug. 3, 1999.

24. Walatka, Pamela P.; Buning, Pieter G.; Pierce, Larry; and Elson, Patricia A.:PLOT3D User's Manual.

NASA TM-101067, 1990.

25. Tecplot Version 7.5 User's Manual. Amtec Eng. Inc., 1988{1998.

26. FIELDVIEW Version 6 User's Guide. Intelligent Light, May 1999.

27. Lamar, John E.; Brandon, Jay; Stacy, Kathryn; Johnson, Thomas D., Jr.; Severance, Kurt; and

Childers, Brooks A.: Leading-Edge Vortex-System Details Obtained on F-106B Aircraft Using a

Rotating Vapor Screen and Surface Techniques. NASA TP-3374, 1993.

28. Bowen, Susan S.: Computer Vision as a Tool for Image Synthesis. International Symposium on

Electronic Imagining, 1995.

29. Wolf, Paul R.: Elements of Photogrammetry. McGraw-Hill, Inc., 1974.

30. Stacy, Kathryn: Computer Graphics|Custom Application: ILLUME. http:==dval-www:larc:nasa:gov=

DVAL=Capabilities=Cg=index7:html Accessed Jan. 4, 2001.

31. Stacy, Kathryn; Severance, Kurt; and Childers, Brooks A.: Computer-Aided Light-Sheet Flow

Visualization Using Photogrammetry. NASA TP-3416, 1994.

32. Stacy, Kathryn: Spotlight Application|Image Reconstruction. http:==dval-www:larc:nasa:gov=DVAL=

Spotlight=index6:html Accessed Jan. 4, 2001.

33. Bingel, Bradford; and Hammond, Dana: Transferable Output ASCII Data (TOAD) File Format

Description. NASA CR-178361, 1987.

34. Bingel, Bradford D.; Shea, Anne L.; and Ho
er, Alicia S.: Transferable Output ASCII Data (TOAD)

Editor Version 1.0 User's Guide. NASA CR-187507, 1991.

35. Reisel, Joseph F.: The F16XL Database. http:=dval-www:larc:nasa:gov=optics=f16xl delivery=

index:html Accessed Jan. 4, 2001.

36. Cronin, Catherine K.: Program for Information Storage and Management|F16XL High Lift Pro ject

(CAWAP). http:==f16xl-prism:larc:nasa:gov : =cgi-bin=webprism:cgi?f16xl@remote-dbsrv1 Accessed

Jan. 4, 2001.

37. Baals, Donald D.; and Corliss, William R.: Wind Tunnels of NASA. NASA SP-440, 1981.

38. Sellers, William L., III; and Kjelgaard, Scott O.: The Basic Aerodynamics Research Tunnel|A Facility

Dedicated to Code Validation. AIAA-88-1997, May 1988.

34
Table 1. Planned F-16XL CAWAP Data Comparisons

Item Data comparison Data source

1 On and o surface 
ow

Tufts Flight

Static pressures in a row Flight

Vapor screen Flight

Particle traces CFD

2 Surface 
ow and pressure

Oil 
ow Wind tunnel

Pressure sensitive paint Wind tunnel

3 Surface pressure

Pressure sensitive paint Wind tunnel

Calculations CFD

4 Vortex core location

Vapor screen Wind tunnel

Particle traces CFD

5 Vortex core location

Vapor screen Flight, wind tunnel

Particle traces CFD

6 Surface 
ow

Oil 
ow Flight

Tufts Flight

Liquid crystals Flight

7 Surface 
ow

Oil 
ow Flight

Tufts Flight

Particle traces CFD

8 Surface 
ow

Oil 
ow Flight, wind tunnel

Particle traces CFD

9 Surface pressure

Pressure sensitive paint Wind tunnel

Static pressures in a row Wind tunnel

10 O surface 
ow

Vapor screen Flight, wind tunnel

Five-hole probe Wind tunnel

11 On and o surface 
ow

Oil 
ow Flight

Vapor screen Flight

Particle traces CFD

12 On and o surface 
ow

Oil 
ow Flight, wind tunnel

Vapor screen Flight, wind tunnel

Five-hole probe Wind tunnel

13 Surface 
ow

Oil 
ow Flight

Tufts Flight

Liquid crystals Flight

Propylene-glycol-methyl-ether traces Flight

35
Table 1. Concluded

Item Data comparison Data source

14 Surface 
ow

Oil 
ow Fli ght

Tufts Flight

Particle traces CFD

Propylene -glycol-methyl-ether traces Flight

15 Vortex core location

Vapor screen Wind tunnel

Dye traces Water tunnel

Particle traces CFD

16 Surface pressure contours

Static pressures Flight, wind tunnel

Calculations CFD

17 Surface pressure

Pressure sensitive paint Flight

Static pressures in a row Flight, wind tunnel

18 Surface pressure

Pressure sensitive paint Flight

Static pressures in a row Flight, CFD

Static-pressure surfaces CFD

19 On and o surface 
ow

Tufts Flight

Particle traces CFD

Stagnation pressure contours CFD

20 Boundary-layer prole

Rake Flight

Velocities CFD

36
Table 2. Actual F-16XL CAWAP Data Comparisons

Item

(a) Data comparison Data source

3 Surface pressure contours

Pressure sensitive paint Wind tunnel

Calculations CFD

6 Surface 
ow

Oil 
ow Flight

Tufts Flight

Liquid crystals Flight

12a On and o surface 
ow

Oil 
ow Wind tunnel

Vapor screen Wind tunnel

16a Surface pressure contours

Static pressures Flight

Calculations CFD

17a Surface pressure

Static pressures in a row Flight, wind tunnel

18a Surface pressure

Static pressures in a row Flight, CFD

Static-pressure surfaces CFD

19 On and o surface 
ow

Tufts Flight

Particle traces CFD

Stagnation pressure contours CFD

20 Boundary-layer prole

Rake Flight

Velocities CFD a\a" in item number signies reduction in number of items being compared.

37
Table 3. Airplane Specications

Aspect ratio, A . . . . . . . . . . . . . . . . . . . . . . 1.75

Wingspan, b, ft . . . . . . . . . . . . . . . . . . . . . . 32.4

Wing area, ft2 ...................... 646.37

Wing surface area, Sre f, ft2 . . . . . . . . . . . . . . . . . . 600

Reference wing chord, c, ft . . . . . . . . . . . . . . . . . 24.7

Theoretical root chord, cr, ft . . . . . . . . . . . . . . . . . 41.75

Height, ft . . . . . . . . . . . . . . . . . . . . . . . . 17.606

Length, ft ........................ 54.155

Typical takeo weight, lb . . . . . . . . . . . . . . . . . 35 000

Engine . . . . . . . . . . . . . . Pratt & Whitney F100-PW-200

Maximum thrust, lb . . . . . . . . . . . . . . . . . . 23 830

38
Table 4. Requested, Actual, and Nominal Aircraft Pressure Port Locations

Port Requested Actual Nominal Port type

(a) FS BL FS BL FS BL (b)

1 160.75 36.00 160.95 35.88 161.10 36.00 Flush, US

2 161.25 36.00 161.45 35.89 161.59 36.00 Flush, US

3 165.50 36.00 165.63 35.88 165.77 36.00 Flush, US

4 166.00 36.00 166.12 35.89 166.24 36.00 Flush, US

5 169.00 36.00 169.12 35.90 169.24 36.00 Flush, US

6 170.00 36.00 170.08 35.90 170.20 36.00 Flush, US

7 171.00 36.00 171.06 35.91 171.16 36.00 Flush, US

8 172.00 36.00 172.06 35.90 172.18 36.00 Flush, US

9 176.00 36.00 176.03 35.93 176.12 36.00 Flush, US

10 185.00 36.00 184.94 35.95 185.00 36.00 Flush, US

11 162.00 37.00 162.19 36.87 162.35 37.00 Flush, US

12 162.50 37.00 162.69 36.87 162.85 37.00 Flush, US

13 163.25 38.00 163.46 37.87 163.61 38.00 Flush, US

14 163.75 38.00 163.93 37.89 164.07 38.00 Flush, US

15 166.00 38.00 167.14 37.88 167.28 38.00 Flush, US

16 169.00 38.00 169.14 37.88 169.29 38.00 Flush, US

17 170.00 38.00 170.13 37.88 170.27 38.00 Flush, US

18 171.00 38.00 171.12 37.88 171.26 38.00 Flush, US

19 172.00 38.00 172.11 37.89 172.24 38.00 Flush, US

20 176.00 38.00 176.08 37.89 176.21 38.00 Flush, US

21 185.00 38.00 184.97 37.90 185.07 38.00 Flush, US

22 164.50 39.00 164.73 38.87 164.88 39.00 Flush, US

23 165.00 39.00 165.23 38.87 165.39 39.00 Flush, US

24 165.50 40.00 165.72 39.87 165.87 40.00 Flush, US

25 166.00 40.00 166.22 39.86 166.38 40.00 Flush, US

26 169.00 40.00 169.18 39.87 169.33 40.00 Flush, US

27 170.00 40.00 170.17 39.87 170.31 40.00 Flush, US

28 171.00 40.00 171.15 39.87 171.29 40.00 Flush, US

29 172.00 40.00 172.14 39.87 172.28 40.00 Flush, US

30 176.00 40.00 176.12 39.87 176.26 40.00 Flush, US

31 180.00 40.00 180.06 39.87 180.20 40.00 Flush, US

32 185.00 40.00 185.02 39.88 185.15 40.00 Flush, US

129 202.50 40.00 202.50 39.78 202.50 40.00 Belt

130 215.00 40.00 215.00 39.84 215.00 40.00 Belt

131 230.00 40.00 230.00 39.78 230.00 40.00 Belt

132 240.00 40.00 Does not exist

133 250.00 40.00 250.00 39.48 250.00 40.00 Belt

134 257.50 40.00 257.50 39.46 257.50 40.00 Belt

135 282.50 40.00 282.50 39.60 282.50 40.00 Belt

136 300.00 40.00 300.00 39.66 300.00 40.00 Belt

137 337.50 40.00 337.50 39.76 337.50 40.00 Belt

138 375.00 40.00 375.00 39.86 375.00 40.00 Belt

139 390.00 40.00 390.00 40.07 390.00 40.00 Belt

140 407.50 40.00 407.50 40.10 407.50 40.00 Belt

141 425.00 40.00 425.00 40.15 425.00 40.00 Belt

142 437.50 40.00 437.50 40.09 437.50 40.00 Belt

143 450.00 40.00 450.00 40.14 450.00 40.00 Belt

a Deleted port.

b LE is leading edge; LS is lower surface ; US is upper surface.

39
Table 4. Continued

Port Requested Actual Nominal Port type

(a) FS BL FS BL FS BL (b)

144 462.50 40.00 462.50 40.18 462.50 40.00 Belt

145 183.00 55.00 183.00 54.09 183.00 55.00 Belt

146 184.00 55.00 184.00 54.17 184.00 55.00 Belt

147 185.00 55.00 185.00 54.24 185.00 55.00 Belt

148 187.50 55.00 187.50 54.37 187.50 55.00 Belt

149 190.00 55.00 190.00 54.45 190.00 55.00 Belt

150 192.50 55.00 192.50 54.55 192.50 55.00 Belt

151 195.00 55.00 195.00 54.66 195.00 55.00 Belt

152 197.50 55.00 197.50 54.73 197.50 55.00 Belt

153 200.00 55.00 200.00 54.75 200.00 55.00 Belt

154 202.50 55.00 202.50 54.84 202.50 55.00 Belt

155 205.00 55.00 205.00 54.97 205.00 55.00 Belt

156 210.00 55.00 210.00 55.06 210.00 55.00 Belt

157 215.00 55.00 215.00 55.14 215.00 55.00 Belt

158 220.00 55.00 220.00 55.21 220.00 55.00 Belt

159 230.00 55.00 230.00 55.26 230.00 55.00 Belt

160 240.00 55.00 240.00 55.64 240.00 55.00 Belt

161 250.00 55.00 250.00 55.72 250.00 55.00 Belt

162 257.50 55.00 257.50 55.75 257.50 55.00 Belt

163 265.00 55.00 265.00 55.81 265.00 55.00 Belt

164 282.50 55.00 282.50 55.86 282.50 55.00 Belt

165 300.00 55.00 300.00 55.97 300.00 55.00 Belt

166 337.50 55.00 337.50 56.06 337.50 55.00 Belt

167 375.00 55.00 375.00 56.07 375.00 55.00 Belt

168 390.00 55.00 390.00 56.16 390.00 55.00 Belt

169 407.50 55.00 407.50 56.16 407.50 55.00 Belt

170 425.00 55.00 425.00 56.18 425.00 55.00 Belt

171 437.50 55.00 Does not exist

172 450.00 55.00 Does not exist

173 462.50 55.00 Does not exist

174 465.00 55.00 Does not exist

182 500.00 55.00 Does not exist

33 208.51 70.00 206.38 69.12 209.56 70.00 Flush, LE

36 210.00 70.00 207.82 69.14 210.92 70.00 Flush LS

37 210.50 70.00 208.33 69.15 211.42 70.00 Flush, LS

34 210.75 70.00 208.58 69.11 211.77 70.00 Flush, LS

38 211.00 70.00 208.81 69.16 211.86 70.00 Flush, LS

35 211.50 70.00 209.31 69.11 212.49 70.00 Flush, LS

39 212.00 70.00 209.76 69.14 212.85 70.00 Flush, LS

40 213.00 70.00 210.82 69.18 213.99 70.00 Flush, LS

41 214.00 70.00 211.84 69.09 215.07 70.00 Flush, LS

42 215.00 70.00 212.81 69.06 216.08 70.00 Flush, LS

43 216.00 70.00 213.84 69.06 217.12 70.00 Flush, LS

44 209.00 70.00 207.10 69.11 210.30 70.00 Flush, US

45 209.50 70.00 207.37 69.13 210.52 70.00 Flush, US

46 210.00 70.00 207.84 69.13 210.99 70.00 Flush, US

47 210.50 70.00 208.31 69.15 211.39 70.00 Flush, US

a Deleted port.

b LE is leading edge; LS is lower surface ; US is upper surface.

40
Table 4. Continued

Port Requested Actual Nominal Port type

(a) FS BL FS BL FS BL (b)

48 211.00 70.00 208.82 69.15 211.91 70.00 Flush, US

49 212.00 70.00 209.77 69.16 212.80 70.00 Flush, US

50 213.00 70.00 210.79 69.20 213.72 70.00 Flush, US

51 214.00 70.00 211.79 69.23 214.64 70.00 Flush, US

52 215.00 70.00 212.76 69.24 215.57 70.00 Flush, US

53 216.00 70.00 213.78 69.26 216.54 70.00 Flush, US

183 220.00 70.00 220.50 67.69 220.50 70.00 Belt

184 222.50 70.00 222.50 68.07 222.50 70.00 Belt

185 225.00 70.00 225.00 68.13 225.00 70.00 Belt

186 227.50 70.00 227.50 68.22 227.50 70.00 Belt

187 230.00 70.00 230.00 68.32 230.00 70.00 Belt

188 232.50 70.00 232.50 68.42 232.50 70.00 Belt

189 235.00 70.00 235.00 68.50 235.00 70.00 Belt

190 237.50 70.00 237.50 68.60 237.50 70.00 Belt

191 240.00 70.00 240.00 68.70 240.00 70.00 Belt

192 245.00 70.00 245.00 68.85 245.00 70.00 Belt

193 250.00 70.00 250.00 68.99 250.00 70.00 Belt

194 255.00 70.00 255.00 69.02 255.00 70.00 Belt

195 257.50 70.00 257.50 69.11 257.50 70.00 Belt

196 265.00 70.00 265.00 69.22 265.00 70.00 Belt

197 270.00 70.00 270.00 69.27 270.00 70.00 Belt

198 275.00 70.00 275.00 69.30 275.00 70.00 Belt

199 282.50 70.00 282.50 69.40 282.50 70.00 Belt

200 290.00 70.00 290.00 69.50 290.00 70.00 Belt

201 295.00 70.00 295.00 69.65 295.00 70.00 Belt

202 300.00 70.00 300.00 69.74 300.00 70.00 Belt

203 310.00 70.00 310.00 69.85 310.00 70.00 Belt

204 320.00 70.00 320.00 69.96 320.00 70.00 Belt

205 330.00 70.00 330.00 70.00 330.00 70.00 Belt

206 337.50 70.00 337.50 70.15 337.50 70.00 Belt

207 350.00 70.00 350.00 70.18 350.00 70.00 Belt

208 360.00 70.00 360.00 70.42 360.00 70.00 Belt

209 375.00 70.00 375.00 70.48 375.00 70.00 Belt

210 390.00 70.00 390.00 70.53 390.00 70.00 Belt

211 407.50 70.00 407.50 70.60 407.50 70.00 Belt

212 425.00 70.00 425.00 70.60 425.00 70.00 Belt

213 437.50 70.00 437.50 70.72 437.50 70.00 Belt

214 450.00 70.00 450.00 70.00 450.00 70.00 Belt

215 462.50 70.00 462.50 70.00 462.50 70.00 Belt

216 465.00 70.00 464.50 70.00 464.50 70.00 Belt

217 470.00 70.00 Does not exist

218 472.50 70.00 Does not exist

219 475.00 70.00 475.00 69.69 475.00 70.00 Belt

220 477.50 70.00 477.50 69.76 477.50 70.00 Belt

221 480.00 70.00 480.00 69.83 480.00 70.00 Belt

222 485.00 70.00 485.00 69.88 485.00 70.00 Belt

223 492.50 70.00 492.50 69.98 492.50 70.00 Belt

a Deleted port.

b LE is leading edge; LS is lower surface ; US is upper surface.

41
Table 4. Continued

Port Requested Actual Nominal Port type

(a) FS BL FS BL FS BL (b)

224 500.00 70.00 500.00 70.07 500.00 70.00 Belt

54 235.50 80.00 232.85 78.97 236.57 80.00 Flush, LE

64 243.00 80.00 233.10 79.02 236.69 80.00 Flush, LS

55 236.00 80.00 233.40 78.92 237.26 80.00 Flush, LS

56 236.50 80.00 233.90 78.96 237.65 80.00 Flush, LS

57 237.00 80.00 234.39 78.97 238.10 80.00 Flush, LS

58 237.50 80.00 234.89 79.01 238.49 80.00 Flush, LS

59 238.00 80.00 235.38 79.02 238.95 80.00 Flush, LS

60 239.00 80.00 236.39 79.05 239.88 80.00 Flush, LS

61 240.00 80.00 237.38 79.07 240.79 80.00 Flush, LS

62 241.00 80.00 238.39 79.08 241.77 80.00 Flush, LS

63 235.70 80.00 239.37 79.11 242.66 80.00 Flush, LS

66 236.70 80.00 232.96 78.97 236.71 80.00 Flush, US

67 237.15 80.00 234.52 79.03 238.08 80.00 Flush, US

65 236.00 80.00 234.77 79.07 238.22 80.00 Flush, US

68 237.50 80.00 234.85 79.03 238.41 80.00 Flush, US

69 238.45 80.00 235.79 79.04 239.31 80.00 Flush, US

71 240.00 80.00 237.26 79.09 240.64 80.00 Flush, US

72 241.00 80.00 238.28 79.12 241.57 80.00 Flush, US

73 242.00 80.00 239.27 79.11 242.56 80.00 Flush, US

74 243.00 80.00 240.24 79.13 243.48 80.00 Flush, US

70 244.00 80.00 241.22 79.14 244.41 80.00 Flush, US

225 250.00 80.00 250.00 78.99 250.00 80.00 Belt

226 257.50 80.00 257.50 79.19 257.50 80.00 Belt

227 270.00 80.00 270.00 79.42 270.00 80.00 Belt

228 282.50 80.00 282.50 79.55 282.50 80.00 Belt

229 300.00 80.00 300.00 79.69 300.00 80.00 Belt

230 320.00 80.00 320.00 79.88 320.00 80.00 Belt

231 337.50 80.00 337.50 79.99 337.50 80.00 Belt

232 350.00 80.00 350.00 80.06 350.00 80.00 Belt

233 375.00 80.00 375.00 80.00 375.00 80.00 Belt

175 470.00 55.00 470.00 80.00 470.00 80.00 Belt

176 472.50 55.00 472.50 80.00 472.50 80.00 Belt

177 475.00 55.00 475.00 80.00 475.00 80.00 Belt

178 477.50 55.00 477.50 80.00 477.50 80.00 Belt

179 480.00 55.00 480.00 80.00 480.00 80.00 Belt

180 485.00 55.00 485.00 80.00 485.00 80.00 Belt

181 492.50 55.00 492.50 80.00 492.50 80.00 Belt

75 276.41 95.00 274.08 93.99 277.09 95.00 Flush, LE

76 277.00 95.00 274.61 94.06 277.43 95.00 Flush, LS

77 277.50 95.00 275.10 94.11 277.76 95.00 Flush, LS

78 278.00 95.00 275.60 94.18 278.08 95.00 Flush, LS

79 278.50 95.00 276.09 94.20 278.50 95.00 Flush, LS

80 279.00 95.00 276.58 94.26 278.84 95.00 Flush, LS

81 280.00 95.00 277.55 94.33 279.61 95.00 Flush, LS

82 281.00 95.00 278.54 94.40 280.42 95.00 Flush, LS

83 282.00 95.00 279.53 94.45 281.24 95.00 Flush, LS

a Deleted port.

b LE is leading edge; LS is lower surface ; US is upper surface.

42
Table 4. Continued

Port Requested Actual Nominal Port type

(a) FS BL FS BL FS BL (b)

84 282.50 95.00 280.02 94.48 281.67 95.00 Flush, LS

85 283.00 95.00 280.53 94.51 282.09 95.00 Flush, LS

86 277.00 95.00 274.60 94.02 277.51 95.00 Flush, US

87 277.50 95.00 275.08 94.05 277.91 95.00 Flush, US

88 278.00 95.00 275.56 94.07 278.31 95.00 Flush, US

89 278.50 95.00 276.05 94.10 278.74 95.00 Flush, US

90 279.00 95.00 276.53 94.12 279.17 95.00 Flush, US

91 280.00 95.00 277.51 94.14 280.07 95.00 Flush, US

92 281.00 95.00 278.48 94.17 280.96 95.00 Flush, US

93 282.00 95.00 279.44 94.18 281.87 95.00 Flush, US

94 282.50 95.00 279.94 94.19 282.36 95.00 Flush, US

95 283.00 95.00 280.42 94.19 282.83 95.00 Flush, US

234 287.50 95.00 Does not exist

235 290.00 95.00 Does not exist

236 292.50 95.00 292.50 92.76 292.50 95.00 Belt

237 295.00 95.00 295.00 92.85 295.00 95.00 Belt

238 297.50 95.00 297.50 92.95 297.50 95.00 Belt

239 300.00 95.00 300.00 93.08 300.00 95.00 Belt

240 302.50 95.00 302.50 93.18 302.50 95.00 Belt

241 305.00 95.00 305.00 93.30 305.00 95.00 Belt

242 307.50 95.00 307.50 93.39 307.50 95.00 Belt

243 310.00 95.00 310.00 93.49 310.00 95.00 Belt

244 315.00 95.00 315.00 93.63 315.00 95.00 Belt

245 320.00 95.00 320.00 93.76 320.00 95.00 Belt

246 325.00 95.00 325.00 93.85 325.00 95.00 Belt

247 330.00 95.00 330.00 93.87 330.00 95.00 Belt

248 337.50 95.00 337.50 94.00 337.50 95.00 Belt

249 350.00 95.00 350.00 94.14 350.00 95.00 Belt

250 360.00 95.00 360.00 94.18 360.00 95.00 Belt

251 375.00 95.00 375.00 94.29 375.00 95.00 Belt

252 390.00 95.00 390.00 94.41 390.00 95.00 Belt

253 407.50 95.00 407.50 94.48 407.50 95.00 Belt

254 425.00 95.00 425.00 94.61 425.00 95.00 Belt

255 437.50 95.00 437.50 95.00 437.50 95.00 Belt

256 450.00 95.00 450.00 95.00 450.00 95.00 Belt

257 462.50 95.00 462.50 95.00 462.50 95.00 Belt

258 465.00 95.00 465.00 95.00 465.00 95.00 Belt

259 470.00 95.00 470.00 94.82 470.00 95.00 Belt

260 472.50 95.00 472.50 94.87 472.50 95.00 Belt

261 475.00 95.00 475.00 94.92 475.00 95.00 Belt

262 477.50 95.00 477.50 94.96 477.50 95.00 Belt

263 480.00 95.00 480.00 95.05 480.00 95.00 Belt

264 485.00 95.00 485.00 95.15 485.00 95.00 Belt

265 492.50 95.00 492.50 95.22 492.50 95.00 Belt

266 500.00 95.00 500.00 95.34 500.00 95.00 Belt

96 304.15 105.00 301.61 104.04 304.09 105.00 Flush, LE

103 304.30 105.00 301.81 104.01 304.36 105.00 Flush, LS

a Deleted port.

b LE is leading edge; LS is lower surface ; US is upper surface.

43
Table 4. Continued

Port Requested Actual Nominal Port type

(a) FS BL FS BL FS BL (b)

97 304.50 105.00 302.13 103.98 304.76 105.00 Flush, LS

98 305.00 105.00 302.61 103.98 305.25 105.00 Flush, LS

99 305.50 105.00 303.14 104.02 305.65 105.00 Flush, LS

100 305.70 105.00 303.35 104.01 305.90 105.00 Flush, LS

101 306.50 105.00 304.12 104.04 306.57 105.00 Flush, LS

102 307.00 105.00 304.63 104.04 307.06 105.00 Flush, LS

104 308.00 105.00 305.62 104.05 308.01 105.00 Flush, LS

105 308.50 105.00 306.12 104.05 308.52 105.00 Flush, LS

107 309.50 105.00 307.13 104.08 309.42 105.00 Flush, LS

108 310.00 105.00 307.65 104.11 309.88 105.00 Flush, LS

106 310.50 105.00 308.12 104.10 310.36 105.00 Flush, LS

109 311.00 105.00 308.64 104.12 310.84 105.00 Flush, LS

110 312.00 105.00 309.62 104.12 311.77 105.00 Flush, LS

111 313.00 105.00 310.65 104.15 312.73 105.00 Flush, LS

112 314.00 105.00 311.66 104.15 313.74 105.00 Flush, LS

116 304.30 105.00 301.83 104.05 304.26 105.00 Flush, US

113 304.50 105.00 302.14 104.06 304.56 105.00 Flush, US

114 305.00 105.00 302.62 104.09 304.96 105.00 Flush, US

115 305.30 105.00 302.94 104.13 305.16 105.00 Flush, US

117 306.70 105.00 304.28 104.14 306.45 105.00 Flush, US

118 307.00 105.00 304.54 104.15 306.70 105.00 Flush, US

119 307.50 105.00 305.02 104.16 307.13 105.00 Flush, US

120 308.00 105.00 305.51 104.17 307.58 105.00 Flush, US

121 308.50 105.00 306.00 104.18 308.05 105.00 Flush, US

122 309.00 105.00 306.49 104.20 308.48 105.00 Flush, US

123 309.50 105.00 306.99 104.21 308.95 105.00 Flush, US

124 310.00 105.00 307.48 104.22 309.41 105.00 Flush, US

125 311.00 105.00 308.46 104.26 310.29 105.00 Flush, US

126 312.00 105.00 309.45 104.28 311.21 105.00 Flush, US

127 313.00 105.00 310.43 104.30 312.13 105.00 Flush, US

128 314.00 105.00 311.41 104.32 313.05 105.00 Flush, US

267 317.50 105.00 Does not exist

268 320.00 105.00 320.00 102.64 320.00 105.00 Belt

269 322.50 105.00 322.50 102.70 322.50 105.00 Belt

270 325.00 105.00 325.00 102.81 325.00 105.00 Belt

271 327.50 105.00 327.50 102.93 327.50 105.00 Belt

272 330.00 105.00 330.00 103.04 330.00 105.00 Belt

273 332.50 105.00 332.50 103.11 332.50 105.00 Belt

274 335.00 105.00 335.00 103.19 335.00 105.00 Belt

275 337.50 105.00 337.50 103.26 337.50 105.00 Belt

276 340.00 105.00 340.00 103.37 340.00 105.00 Belt

277 342.50 105.00 342.50 103.47 342.50 105.00 Belt

278 345.00 105.00 345.00 103.55 345.00 105.00 Belt

279 347.50 105.00 347.50 103.62 347.50 105.00 Belt

280 350.00 105.00 350.00 103.67 350.00 105.00 Belt

281 355.00 105.00 355.00 103.68 355.00 105.00 Belt

282 360.00 105.00 360.00 103.76 360.00 105.00 Belt

a Deleted port.

b LE is leading edge; LS is lower surface ; US is upper surface.

44
Table 4. Continued

Port Requested Actual Nominal Port type

(a) FS BL FS BL FS BL (b)

283 365.00 105.00 365.00 103.86 365.00 105.00 Belt

284 370.00 105.00 370.00 103.93 370.00 105.00 Belt

285 375.00 105.00 375.00 104.01 375.00 105.00 Belt

286 390.00 105.00 390.00 104.12 390.00 105.00 Belt

287 407.50 105.00 407.50 104.24 407.50 105.00 Belt

288 425.00 105.00 425.00 104.31 425.00 105.00 Belt

289 437.50 105.00 437.50 104.41 437.50 105.00 Belt

290 450.00 105.00 450.00 105.00 450.00 105.00 Belt

291 462.50 105.00 462.50 105.00 462.50 105.00 Belt

292 465.00 105.00 465.00 105.00 465.00 105.00 Belt

293 470.00 105.00 470.00 104.63 470.00 105.00 Belt

294 472.50 105.00 472.50 104.75 472.50 105.00 Belt

295 475.00 105.00 475.00 104.84 475.00 105.00 Belt

296 477.50 105.00 477.50 104.92 477.50 105.00 Belt

297 480.00 105.00 480.00 105.00 480.00 105.00 Belt

298 485.00 105.00 485.00 105.09 485.00 105.00 Belt

299 492.50 105.00 492.50 105.19 492.50 105.00 Belt

300 500.00 105.00 500.00 105.28 500.00 105.00 Belt

301 370.00 127.50 Does not exist

302 372.50 127.50 372.50 126.71 372.50 127.50 Belt

303 375.00 127.50 375.00 126.87 375.00 127.50 Belt

304 377.50 127.50 377.50 127.03 377.50 127.50 Belt

305 380.00 127.50 380.00 127.18 380.00 127.50 Belt

306 382.50 127.50 382.50 127.30 382.50 127.50 Belt

307 385.00 127.50 385.00 127.37 385.00 127.50 Belt

308 390.00 127.50 390.00 127.52 390.00 127.50 Belt

309 395.00 127.50 395.00 127.66 395.00 127.50 Belt

310 400.00 127.50 400.00 127.70 400.00 127.50 Belt

311 407.50 127.50 407.50 127.80 407.50 127.50 Belt

312 415.00 127.50 415.00 127.88 415.00 127.50 Belt

313 425.00 127.50 425.00 127.94 425.00 127.50 Belt

314 415.00 153.50 415.00 152.86 415.00 153.50 Belt

315 417.50 153.50 417.50 153.02 417.50 153.50 Belt

316 420.00 153.50 420.00 153.11 420.00 153.50 Belt

317 422.50 153.50 422.50 153.24 422.50 153.50 Belt

318 425.00 153.50 425.00 153.35 425.00 153.50 Belt

319 427.50 153.50 427.50 153.47 427.50 153.50 Belt

320 430.00 153.50 430.00 153.50 430.00 153.50 Belt

321 432.50 153.50 432.50 153.57 432.50 153.50 Belt

322 437.50 153.50 Does not exist

323 445.00 153.50 445.00 152.91 445.00 153.50 Belt

324 450.00 153.50 450.00 153.01 450.00 153.50 Belt

325 455.00 153.50 455.00 153.01 455.00 153.50 Belt

326 462.50 153.50 462.50 153.22 462.50 153.50 Belt

327 470.00 153.50 470.00 153.29 470.00 153.50 Belt

328 475.00 153.50 475.00 153.41 475.00 153.50 Belt

329 477.50 153.50 477.50 153.45 477.50 153.50 Belt

a Deleted port.

b LE is leading edge; LS is lower surface ; US is upper surface.

45
Table 4. Concluded

Port Requested Actual Nominal Port type

(a) FS BL FS BL FS BL (b)

330 485.00 153.50 485.00 153.28 485.00 153.50 Belt

331 487.50 153.50 487.50 153.37 487.50 153.50 Belt

332 490.00 153.50 490.00 153.44 490.00 153.50 Belt

333 492.50 153.50 492.50 153.51 492.50 153.50 Belt

334 495.00 153.50 495.00 153.59 495.00 153.50 Belt

335 500.00 153.50 500.00 153.73 500.00 153.50 Belt

336 510.00 153.50 510.00 153.50 510.00 153.50 Belt

337 450.00 184.50 450.00 184.06 450.00 184.50 Belt

338 452.50 184.50 452.50 184.19 452.50 184.50 Belt

339 455.00 184.50 455.00 184.30 455.00 184.50 Belt

340 457.50 184.50 457.50 184.43 457.50 184.50 Belt

341 460.00 184.50 460.00 184.55 460.00 184.50 Belt

342 465.00 184.50 465.00 184.12 465.00 184.50 Belt

343 470.00 184.50 470.00 184.22 470.00 184.50 Belt

344 475.00 184.50 475.00 184.27 475.00 184.50 Belt

345 480.00 184.50 480.00 184.39 480.00 184.50 Belt

346 485.00 184.50 485.00 184.49 485.00 184.50 Belt

347 487.50 184.50 487.50 184.57 487.50 184.50 Belt

348 490.00 184.50 Does not exist

349 495.00 184.50 Does not exist

350 497.50 184.50 Does not exist

351 500.00 184.50 500.00 184.36 500.00 184.50 Belt

352 510.00 184.50 510.00 184.45 510.00 184.50 Belt

353 520.00 184.50 520.00 184.53 520.00 184.50 Belt

a Deleted port.

b LE is leading edge; LS is lower surface ; US is upper surface.

46
Run number for flight number of —

Table 5. F-16XL CAWAP Flight Conditions and Associated Flights and Runs

[PT, Preston tube; RK, boundary-layer rake; HF, hot film; FV, flow visualization; WUT, wind-up turn; PO, pushover; NA, 

βnom,

deg

αnom,

deg

Loadnom,

g units

a131

PT(d)

aFlight pressure data not available.bFlight pressure data available for all or most runs.cFlight pressure data available for some runs.

a132

PT

a133

PT

a134

RK

b135

RK

(d)

b137

RK

(d)

b138

RK, HF

(d)

a139

HF

a141

HF

a142

HF

a143

FV

c144

FV

(d)

a145

FV

(d)

146

FV

a147

FV

Rθ

5 0 1 2b 17b 17b 17b 10b 17b 17b 14b 

3.46 248.27

5 +5 1 2c 17c 17c 17c 10c 17c 17c 14c 

3.46 293.35

5 –5 1 2d 17d 17d 17d 10d 17d 17d 14d 

3.46 215.41

10 0 1 3b 18b 18b 18b 11b 18b 15b 

2.16 196.16

10 +5 1 3c 18c 18c 18c 11c 18c 15c 

2.16 231.78

10 –5 1 3d 18d 18d 18d 11d 18d 

2.16 170.20

13 0 1 4b 19b 12b 19b 19b 19b 16b 1.79 178.5713 +5 1 4c 19c 12c 19c 19c 19c 16c 1.79 210.9913 –5 1 4d 19d 12d 19d 19d 19d 16d 1.79 154.9415 0 1 5b 13b 20b 20b 20b 17b 

1.61 169.35

15 +5 1 5c 13c 20c 20c 20c 17c 

1.61 200.11

15 –5 1 5d 13d 20d 20d 20d 17d 

1.61 146.94

5 0 1 12b 12b 12b 12b 7b 12b 12b 8b 

3.26 240.98

5 +5 1 12f 12c 12c 12c 7c 12c 12c 8c 

3.26 284.74

5 –5 1 12g 12d 12d 12d 7d 12d 12d 8d 

3.26 209.09

10 0 1 13b 13b 13b 13b 9b 3c 3b 13b 13b 9b 

2.03 190.16

10 +5 1 13f 13c 13c 13c 9c 3d 3c 13c 13c 9c 

2.03 224.69

10 –5 1 13g 13d 13d 13d 9d 3e 3d 13d 13d 9d 

2.03 165.00

13 0 1 14b 14b 11b 14b 14b 14b 14b 10b 

1.71 174.53

13 +5 1 14f 14c 11c 14c 14c 14c 14c 10c 

1.71 206.23

13 –5 1 14g 14d 11d 14d 14d 14d 14d 10d 

1.71 151.44

15 0 1 15b 15b 15b 15b 15b 15b 11b 1.56 166.7015 +5 1 15f 15c 15c 15c 15c 15c 11c 1.56 196.9715 –5 1 15g 15d 15d 15d 15d 15d 11d 1.56 144.6420 0 1 16b 16b 16b 16b 8b 16b 16b 1.28 151.0020 +5 1 16c 16c 16c 16c 8c 16c 16c 1.28 178.4220 –5 1 16d 16d 16d 16d 8d 16d 16d 1.28 131.02

5 0 1 7b 6b 7b 7b 7b 4b 7b 7b 5b 3.03 232.335 +5 1 7c 6c 7c 7c 7c 4c 7c 7c 5c 

3.03 274.52

5 –5 1 7d 6d 7d 7d 7d 4d 7d 7d 5d 

3.03 201.58

10 0 1 8b 7b 8b 8b 8b 5b 8b 8b 6b 1.88 183.0010 +5 1 8c 7c 8c 8c 8c 5c 8c 8c 6c 1.88 216.2310 –5 1 8d 7d 8d 8d 8d 5d 8d 8d 6d 1.88 158.79

FC

123456789

101112131415161718192021222324252627282930313233

Mnom

0.560.560.560.350.350.350.290.290.290.260.260.260.610.610.610.380.380.380.320.320.320.290.290.290.240.240.240.710.710.710.440.440.44

hnom,

ft

5 000

5 0005 0005 0005 0005 0005 0005 0005 0005 0005 0005 000

10 00010 00010 00010 00010 00010 00010 00010 00010 00010 00010 00010 00010 00010 00010 00017 50017 50017 50017 50017 50017 500

Rn/ft

× 10–6

d Preston tube or boundary-layer runs.

First choice for flow when pressure file exists.First choice for flow when pressure file does not exist.

47
Run number for flight number of —

βnom,

deg

αnom,

deg

Loadnom,

g units

a131

PT(d)

aFlight pressure data not available.bFlight pressure data available for all or most runs.cFlight pressure data available for some runs.

a132

PT

a133

PT

a134

RK

b135

RK

(d)

b137

RK

(d)

b138

RK, HF

(d)

a139

HF

a141

HF

a142

HF

a143

FV

c144

FV

(d)

a145

FV

(d)

146

FV

a147

FV

Rθ

13 0 1 9b 8b 9b 10b 9b 9b 9b 9b 1.58 167.7713 +5 1 9c 8c 9c 10c 9c 9c 9c 9c 1.58 198.2313 –5 1 9d 8d 9d 10d 9d 9d 9d 9d 1.58 145.5715 0 1 10b 9b 10b 10b 10b 10b 10b 1.45 160.7215 +5 1 10c 9c 10c 10c 10c 10c 10c 1.45 189.9015 –5 1 10d 9d 10d 10d 10d 10d 10d 1.45 139.4520 0 1 11b 10b 11b 11b 11b 6b 11b 11b 7b 1.20 146.2120 +5 1 11c 10c 11c 11c 11c 6c 11c 11c 7c 1.20 172.7620 –5 1 11d 10d 11d 11d 11d 6d 11d 11d 

1.20 126.86

5 0 1 2b 4b 2c 2c 2c 2c 2c 2c 

2.82 224.13

5 +5 1 2c 4c 2d 2d 2d 2d 2d 2d 

2.82 264.83

5 –5 1 2d 4d 2e 2e 2e 2E 2e 2e 

2.82 194.47

10 0 1 3b 5b 3b 3b 3b 3b 1.77 177.5710 +5 1 3c 5c 3c 3c 3c 3c 1.77 209.8110 –5 1 3d 5d 3d 3d 3d 3d 1.77 154.0713 0 1 4b 6b 4b 4b 4b 4b 1.46 161.2713 +5 1 4c 6c 4c 4c 4c 4c 1.46 190.56

FC

3435363738394041424344454647484950

Mnom 

0.370.370.370.340.340.340.280.280.280.810.810.810.510.510.510.420.42

hnom,

ft

17 500

17 50017 50017 50017 50017 50017 50017 50017 50024 00024 00024 00024 00024 00024 00024 00024 000

Rn/ft

× 10–6

51 0.42 24 000 13 –5 1 4d 6d 4d 4d 4d 4d 1.46 139.93

52 0.38 24 000 15 0 1 5b 7b 5b 5b 5b 5b 1.32 153.3453 0.38 24 000 15 +5 1 5c 7c 5c 5c 5c 5c 1.32 181.1954 0.38 24 000 15 –5 1 5d 7d 5d 5d 5d 5d 1.32 133.0555 0.32 24 000 20 0 1 6b 8b 6b 6b 3b 6b 6b 3b 1.11 140.6256 0.32 24 000 20 +5 1 6c 8c 6c 6c 3c 6c 6c 3c 

1.11 166.15

57 0.32 24 000 20 –5 1 6d 8d 6d 6d 3d 6d 6d 3d 

1.11 122.01

58 0.51 5 000 10 2.2 LT 3e 18e 18e 

3.15 236.88

59 0.51 10 000 10 1.8 LT 8b 13e 13e 13e 13e 9e 2.73 220.5360 0.51 17 500 10 1.3 LT 8e 11e 8e 6e 2.18 197.0661 0.42 5 000 13 2.1 LT 4e 19e 12e 19e 19e 

2.60 215.21

62 0.42 10 000 13 1.7 LT 8c 14e 11e 14e 14e 14e 10e 2.25 200.2063 0.42 17 500 13 1.3 LT 10e 9e 9e 9e 1.79 178.5764 0.38 5 000 15 2.1 LT 5e 13e 20e 20e 

2.35 204.60

65 0.38 10 000 15 1.7 LT 8d 15e 15e 15e 15e 11e 

2.03 190.16

66

0.38 17 500 15 1.3 LT 10e 10e 10e 1.62 169.88 

000000000

Table 5. Continued

[PT, Preston tube; RK, boundary-layer rake; HF, hot film; FV, flow visualization; WUT, wind-up turn; PO, pushover; NA, 

d Preston tube or boundary-layer runs.

First choice for flow when pressure file exists.First choice for flow when pressure file does not exist.

48
67 0.60 6 000 4.6 1 6b 2c 2c 13c 2c 12c 

3.60 253.24

68 0.90 19 600 3.7 1 9b 5b 4b 5b 10b 3.60 253.2469 0.95 21 300 3.6 1 10b 6b 6b 2b 

3.60 253.24

70 0.98 22 300 3.6 1 11b 7b 7b 3b 

3.60 253.24

71 1.10 25 800 3.3 1 13b 9b 9b 4b 

3.60 253.24

72 1.30 30 800 3.0 1 11b 10b 10b 8b 

3.60 253.24

73 1.50 34 900 2.7 1 3b 12b 11b 6b 

3.60 253.24

74 0.90 8 700 2.9 1 7b 4b 12b 3b 11b 

5.00 298.45

75 0.90 36 800 6.3 1 4b 13b 12b 5b 

2.00 188.75

76 1.30 20 800 2.3 1 12b 8b 8b 9b 

5.00 298.45

77 1.30 44 500 4.9 1 5b 14b 13b 7b 

2.00 188.75

78 0.60 6 000 12.0 3.7 WUT 6d 2e 2e 13e 2e 12e 

3.60 253.24

79 0.90 19 600 9.3 3.7 WUT 9d 5d 4d 5d 10d 3.60 253.2480 0.95 21 300 8.8 3.5 WUT 10d 6d 6d 2d 3.60 253.2481 0.98 22 300 8.1 3.2 WUT 11d 7d 7d 3d 

3.60 253.24

82 1.10 25 800 7.4 3.0 WUT 13d 9d 9d 4d 

3.60 253.24

83 1.30 30 800 6.5 2.7 WUT 11d 10d 10d 8d 

3.60 253.24

84 1.50 34 900 5.8 2.5 WUT 3d 12d 11d 6d 3.60 253.2485 0.90 8 700 8.4 5.0 WUT 7d 4d 12d 3d 11d 

5.00 298.45

86 0.90 36 800 10.5 2.0 WUT 4d 13d 12d 5d 

2.00 188.75

87 1.30 20 800 5.1 3.0 WUT 12d 8d 8d 9d 

5.00 298.45

88 1.30 44 500 6.8 1.5 WUT 5d 14d 13d 7d 

2.00 188.75

89 0.60 6 000 0.5 PO 6c 2d 2d 13d 2d 12d 

3.60 253.24

90 0.90 19 600 0.5 PO 9c 5c 4c 5c 10c 3.60 253.2491 0.95 21 300 0.5 PO 10c 6c 6c 2c 3.60 253.2492 0.98 22 300 0.5 PO 11c 7c 7c 3c 

3.60 253.24

93 1.10 25 800 0.5 PO 13c 9c 9c 4c 

3.60 253.24

94 1.30 30 800 0.5 PO 11c 10c 10c 8c 

3.60 253.24

95 1.50 34 900 0.5 PO 3c 12c 11c 6c 3.60 253.2496 0.90 8 700 0.5 PO 7c 4c 12c 3c 11c 

5.00 298.45

97 0.90 36 800 0.5 PO 4c 13c 12c 5c 

2.00 188.75

98 1.30 20 800 0.5 PO 12c 8c 8c 9c 

5.00 298.45

99

1.30 44 500 

NANANANANANANANANANANA

0000000000000000000000

00000000000 0.5 PO 5c 14c 13c 7c 

2.00 188.75

Run number for flight number of —

Table 5. Concluded

[PT, Preston tube; RK, boundary-layer rake; HF, hot film; FV, flow visualization; WUT, wind-up turn; PO, pushover; NA, n

βnom,

deg

αnom,

deg

Loadnom,

g units

a131

PT(d)

aFlight pressure data not available.bFlight pressure data available for all or most runs.cFlight pressure data available for some runs.

a132

PT

a133

PT

a134

RK

b135

RK

(d)

b137

RK

(d)

b138

RK, HF

(d)

a139

HF

a141

HF

a142

HF

a143

FV

c144

FV

(d)

a145

FV

(d)

146

FV

a147

FV

Rθ FC Mnom hnom, 

ft

Rn/ft

× 10–6

d Preston tube or boundary-layer runs.

First choice for flow when pressure file exists.First choice for flow when pressure file does not exist.

49
Table 6. F-16XL CAWAP Tuft Video Times for Flights and Runs at Spec ied Flight Conditions

V ideo (yr, 1966) hnom, nom, nom, Loadnom;

Flight Run day:hr:min:sec FC Mnom ft deg deg g units

144 3b 074:09:18:30 46 0.51 24000 10 0 1

144 3c 074:09:19:09 47 0.51 24000 10 5 1

144 3d 074:09:19:28 48 0.51 24000 10 5 1

144 4b 074:09:20:36 49 0.42 24000 13 0 1

144 4c 074:09:21:13 50 0.42 24000 13 5 1

144 4d 074:09:21:31 51 0.42 24000 13 5 1

144 5b 074:09:22:35 52 0.38 24000 15 0 1

144 5c 074:09:23:15 53 0.38 24000 15 5 1

144 5d 074:09:23:35 54 0.38 24000 15 5 1

144 6b 074:09:24:30 55 0.32 24000 21 0 1

144 9b 074:09:37:30 34 0.37 17500 13 0 1

144 9c 074:09:38:11 35 0.37 17500 13 5 1

144 9d 074:09:38:35 36 0.37 17500 13 5 1

144 9e 074:09:39:28 63 0.42 17500 13 0 1

144 10b 074:09:41:17 37 0.34 15700 15 0 1

144 10c 074:09:42:22 38 0.34 17500 15 5 1

144 10d 074:09:42:43 39 0.34 17500 15 5 1

144 10e 074:09:43:23 66 0.38 17500 15 0 1.3

144 11b 074:09:45:06 40 0.28 17500 21 0 1

144 11c 074:09:45:54 41 0.28 17500 21 5 1

144 16b 074:10:04:03 25 0.24 10000 21 0 1

144 16c 074:10:04:43 26 0.24 10000 21 5 1

144 16d 074:10:05:03 27 0.24 10000 21 5 1

145 5b 078:13:26:12 28 0.71 17500 5 0 1

145 6b 078:13:29:09 31 0.44 17500 10 0 1

145 6c 078:13:30:04 32 0.44 17500 10 5 1

145 6d 078:13:30:30 33 0.44 17500 10 5 1

145 6e 078:13:31:44 60 0.51 17500 10 0 1.3

145 9e 078:13:41:36 59 0.51 10000 10 0 1.8

145 10e 078:13:46:19 62 0.42 10000 13 0 1.7

145 11b 078:13:47:53 22 0.29 10000 15 0 1

145 11c 078:13:48:47 23 0.29 10000 15 5 1

145 11d 078:13:49:18 24 0.29 10000 15 5 1

145 16b 078:14:03:44 7 0.29 5000 13 0 1

145 16c 078:14:04:47 8 0.29 5000 13 5 1

145 16d 078:14:05:23 9 0.29 5000 13 5 1

145 17b 078:14:07:01 10 0.26 5000 15 0 1

146 3b 079:13:06:49 74 0.9 8700 2.9 0 1

146 5b 079:13:10:50 68 0.9 19600 3.7 0 1

146 5d 079:13:13:32 79 0.9 19600 9.3 0 3.7

146 6d 079:13:16:17 80 0.95 21300 8.8 0 3.5

146 10b 079:13:32:15 72 1.3 30800 3 0 1

146 11d 079:13:38:09 84 1.5 34900 5.8 0 2.5

151 2c a 100:08:20:15 43 0.81 24000 5 0 1

151 2d a 100:08:20:56 44 0.81 24000 5 5 1

151 2e a 100:08:21:17 45 0.81 24000 5 5 1

151 3b a 100:08:27:27 45 0.32 24000 20 5 1

aFlight where reported video times were 1 hr earlier than other data because chronometers for video tim e code inserters were not updated

when change m ade to Pacic daylight tim e.

50
Table 6. Conc luded

Video (yr, 1966) hnom, nom, nom, Loadnom;

Flight Run day:hr:min:sec FC Mnom ft deg deg g units

151 3c a100:08:27:43 57 0.32 24000 20 5 1

151 5b a100:08:29:36 42 0.28 17500 20 5 1

151 4b a100:08:38:48 29 0.71 17500 5 5 1

151 4c a100:08:39:08 30 0.71 17500 5 5 1

151 6b a

100:08:44:11 13 0.61 10000 5 0 1

151 6c a100:08:44:51 14 0.61 10000 5 5 1

151 6d a100:08:45:09 15 0.61 10000 5 5 1

152 2c 102:09:14:51 67 0.6 6000 4.6 0 1

152 2e 102:09:16:16 78 0.6 6000 12 0 3.7

152 3d 102:09:19:48 85 0.9 8700 8.4 0 5

152 4b 102:09:23:37 69 0.95 21300 3.6 0 1

152 5b 102:09:26:51 70 0.98 22300 3.6 0 1

152 5d 102:09:28:13 81 0.98 22300 8.1 0 3.2

152 6b 102:09:36:15 76 1.3 20800 2.3 0 1

152 6d 102:09:37:40 87 1.3 20800 5.1 0 3

152 7b 102:09:33:54 21 1.1 25800 3.3 0 1

152 7d 102:09:35:03 82 1.1 25800 7.4 0 3

152 8b 102:09:49:31 73 1.5 34900 2.7 0 1

152 9b 102:09:46:09 75 0.9 36800 6.3 0 1

152 9d 102:09:47:30 86 0.9 36800 10.5 0 2

152 11b 102:09:55:40 16 0.38 10000 10 0 1

152 11c 102:09:57:08 17 0.38 10000 10 5 1

152 11d 102:09:56:45 18 0.38 10000 10 5 1

151 6d a100:08:45:09 15 0.61 10000 5 5 1

151 6d a100:08:45:09 15 0.61 10000 5 5 1

151 6d a

100:08:45:09 15 0.61 10000 5 5 1

151 6d a100:08:45:09 15 0.61 10000 5 5 1

151 6d a100:08:45:09 15 0.61 10000 5 5 1

151 6d a

100:08:45:09 15 0.61 10000 5 5 1

151 6d a100:08:45:09 15 0.61 10000 5 5 1

152 12b 102:10:04:06 1 0.56 5000 5 0 1

152 12c 102:10:04:58 3 0.56 5000 5 5 1

152 13b 102:10:07:45 4 0.35 5000 10 0 1

152 13c 102:10:08:26 5 0.35 5000 10 5 1

152 13d 102:10:08:48 6 0.35 5000 10 5 1

152 13e 102:10:10:10 58 0.51 5000 10 0 2.2

152 15c 102:10:15:55 11 0.26 5000 15 5 1

153 2c 107:13:29:56 2 0.56 5000 5 5 1

153 3b 107:13:31:34 61 0.42 5000 13 0 2.1

153 4b 107:13:33:27 12 0.26 5000 15 5 1

153 4c 107:13:35:06 64 0.38 5000 15 0 2.1

153 5b 107:13:39:09 19 0.32 10000 13 0 1

153 5c 107:13:40:01 20 0.32 10000 13 5 1

153 5d 107:13:40:27 21 0.32 10000 13 5 1

153 6b 107:13:41:16 65 0.38 10000 15 0 1.7

153 8b 107:13:54:40 77 1.3 44500 4.9 0 1

153 8d 107:13:55:32 88 1.3 44500 6.8 0 1.5

154 5f 109:10:37:48 83 1.3 30800 6.5 0 2.7

aFlight where reported video times were 1 hr earlier than other data because chronometers for video tim e code inserters were not updated

when change m ade to Pacic daylight tim e.

51
52
Table 8. Actual and Nominal Pressure Port Locations at Airplane Scale for 0.11-Scaled Model

Actual Nominal

Port FS BL FS BL

293 175.09 48.00 175.09 48.00

294 183.93 48.00 183.93 48.00

295 194.98 48.00 194.98 48.00

296 209.33 48.00 209.33 49.00

297 242.10 48.00 242.10 48.00

298 281.87 48.00 281.87 48.00

299 328.62 48.00 328.62 48.00

300 372.07 48.00 372.07 48.00

301 417.35 48.00 417.35 48.00

302 462.27 48.50 462.27 48.00

303 475.53 48.50 475.53 48.00

304 487.68 48.50 487.68 48.00

305 504.24 48.50 504.24 48.00

306 516.02 48.50 516.02 48.00

308 211.73 69.00 211.73 69.00

309 220.66 69.00 220.66 69.00

310 235.75 69.00 235.75 69.00

311 257.01 69.00 257.01 69.00

312 284.12 69.00 284.12 69.00

313 317.08 69.00 317.08 69.00

314 356.51 69.00 356.51 69.00

315 392.55 69.00 392.55 69.00

316 427.35 69.00 427.35 69.00

317 462.78 69.00 462.78 69.00

318 475.41 69.00 475.41 69.00

319 486.50 69.00 486.50 69.00

320 502.52 69.00 502.52 69.00

323 262.86 88.00 262.86 88.00

324 270.21 88.00 270.21 88.00

325 282.64 88.00 282.64 88.00

326 299.89 88.00 299.89 88.00

327 322.21 88.00 322.21 88.00

328 349.35 88.00 349.35 88.00

329 381.56 88.00 381.56 88.00

330 411.23 88.00 411.23 88.00

331 437.10 88.00 437.10 88.00

332 462.97 88.00 462.97 88.00

333 475.14 88.00 475.14 88.00

334 485.80 88.00 485.80 88.00

335 500.76 88.00 500.76 88.00

338 319.37 109.00 319.37 109.00

339 324.79 109.00 324.79 109.00

340 334.27 109.00 334.27 109.00

341 347.24 109.00 347.24 109.00

342 364.26 109.00 364.26 109.00

343 384.77 109.00 384.77 109.00

344 409.15 109.00 409.15 109.00

345 431.60 109.00 431.60 109.00

346 447.46 109.00 447.46 109.00

53
Table 8. Continued

Actual Nominal

Port FS BL FS BL

347 463.33 109.00 463.33 109.00

348 474.75 109.00 474.75 109.00

349 484.81 109.00 484.81 109.00

350 498.93 109.00 498.93 109.00

353 378.52 131.00 378.52 131.00

354 382.40 131.00 382.40 131.00

355 388.70 131.00 388.70 131.00

356 397.68 131.00 397.68 131.00

357 409.21 131.00 409.21 131.00

358 423.29 131.00 423.29 131.00

359 439.91 131.00 439.91 131.00

360 455.19 131.00 455.19 131.00

361 468.59 131.00 468.59 131.00

362 478.65 131.00 478.65 131.00

363 487.90 131.00 487.90 131.00

364 500.77 131.00 500.77 131.00

368 401.86 142.00 401.86 142.00

369 406.59 142.00 406.59 142.00

370 416.29 142.00 416.29 142.00

371 424.26 142.00 424.26 142.00

372 433.96 142.00 433.96 142.00

373 446.89 142.00 446.89 142.00

374 459.82 142.00 459.82 142.00

375 472.75 142.00 472.75 142.00

376 482.34 142.00 482.34 142.00

377 491.23 142.00 491.23 142.00

378 503.58 142.00 503.58 142.00

383 419.11 157.00 419.11 157.00

384 422.75 157.00 422.75 157.00

385 430.72 157.00 430.72 157.00

386 437.18 157.00 437.18 157.00

387 445.06 157.00 445.06 157.00

388 456.26 157.00 456.26 157.00

389 467.27 157.00 467.27 157.00

390 478.37 157.00 478.37 157.00

391 487.46 157.00 487.46 157.00

392 495.84 157.00 495.84 157.00

393 506.84 157.00 506.84 157.00

398 435.76 172.00 435.76 172.00

399 438.79 172.00 438.79 172.00

400 445.01 172.00 445.01 172.00

401 450.11 172.00 450.11 172.00

402 456.25 172.00 456.25 172.00

403 465.50 172.00 465.50 172.00

404 474.84 172.00 474.84 172.00

405 484.09 172.00 484.09 172.00

406 492.56 172.00 492.56 172.00

407 500.34 172.00 500.34 172.00

408 511.32 172.00 511.32 172.00

54
Table 8. Continued

Actual Nominal

Port FS BL FS BL

413 452.76 187.00 452.76 187.00

414 454.99 187.00 454.99 187.00

415 459.30 187.00 459.30 187.00

416 462.97 187.00 462.97 187.00

417 467.43 187.00 467.43 187.00

418 474.92 187.00 474.92 187.00

419 482.32 187.00 482.32 187.00

420 489.81 187.00 489.81 187.00

421 497.65 187.00 497.65 187.00

422 504.98 187.00 504.98 187.00

423 515.20 187.00 515.20 187.00

458 196.19 48.00 196.19 48.00

459 204.48 48.00 204.48 48.00

461 228.31 48.00 228.31 48.00

463 296.35 48.00 296.35 48.00

465 380.97 48.00 380.97 48.00

467 465.60 47.50 465.60 48.00

468 478.03 47.50 478.03 48.00

469 489.43 47.50 489.43 48.00

470 504.97 47.50 504.97 48.00

473 211.73 69.00 211.73 69.00

474 220.66 69.00 220.66 69.00

476 257.01 69.00 257.01 69.00

478 317.08 69.00 317.08 69.00

480 392.55 69.00 392.55 69.00

482 462.78 69.00 462.78 69.00

483 475.41 69.00 475.41 69.00

484 486.50 69.00 486.50 69.00

485 502.52 69.00 502.52 69.00

488 262.86 88.00 262.86 88.00

489 270.21 88.00 270.21 88.00

491 299.89 88.00 299.89 88.00

493 349.35 88.00 349.35 88.00

495 411.23 88.00 411.23 88.00

497 462.97 88.00 462.97 88.00

498 475.14 88.00 475.14 88.00

499 485.80 88.00 485.80 88.00

500 500.76 88.00 500.76 88.00

503 319.37 109.00 319.37 109.00

504 324.79 109.00 324.79 109.00

506 347.24 109.00 347.24 109.00

508 384.77 109.00 384.77 109.00

510 431.60 109.00 431.60 109.00

512 463.33 109.00 463.33 109.00

513 474.75 109.00 474.75 109.00

514 484.81 109.00 484.81 109.00

515 498.93 109.00 498.93 109.00

518 378.52 131.00 378.52 131.00

519 382.40 131.00 382.40 131.00

55
Table 8. Concluded

Actual Nominal

Port FS BL FS BL

521 397.68 131.00 397.68 131.00

523 423.29 131.00 423.29 131.00

525 455.19 131.00 455.19 131.00

526 468.59 131.00 468.59 131.00

527 478.65 131.00 478.65 131.00

528 487.90 131.00 487.90 131.00

529 500.77 131.00 500.77 131.00

533 401.86 142.00 401.86 142.00

534 406.59 142.00 406.59 142.00

536 424.26 142.00 424.26 142.00

537 433.96 142.00 433.96 142.00

538 446.89 142.00 446.89 142.00

540 472.75 142.00 472.75 142.00

541 482.34 142.00 482.34 142.00

542 491.23 142.00 491.23 142.00

543 503.58 142.00 503.58 142.00

548 419.11 157.00 419.11 157.00

549 422.75 157.00 422.75 157.00

551 437.18 157.00 437.18 157.00

552 445.06 157.00 445.06 157.00

553 456.26 157.00 456.26 157.00

555 478.37 157.00 478.37 157.00

556 487.46 157.00 487.46 157.00

557 495.84 157.00 495.84 157.00

558 507.45 157.00 507.45 157.00

563 435.76 172.00 435.76 172.00

564 438.79 172.00 438.79 172.00

566 450.11 172.00 450.11 172.00

567 456.25 172.00 456.25 172.00

568 465.50 172.00 465.50 172.00

570 484.09 172.00 484.09 172.00

571 492.56 172.00 492.56 172.00

572 500.34 172.00 500.34 172.00

573 511.32 172.00 511.32 172.00

578 452.76 187.00 452.76 187.00

579 454.99 187.00 454.99 187.00

581 462.97 187.00 462.97 187.00

582 467.43 187.00 467.43 187.00

583 474.92 187.00 474.92 187.00

585 489.81 187.00 489.81 187.00

586 497.65 187.00 497.65 187.00

587 504.98 187.00 504.98 187.00

588 515.20 187.00 515.20 187.00

56
Table 9. Actual and Nominal Pressure Port Locations at Airplane Scale for 0.18-Scaled Model

Actual Nominal

Port FS BL FS BL

1 184.87 55.49 185.00 55.00

2 201.55 55.49 201.55 55.00

3 215.17 55.49 215.00 55.00

4 250.20 55.49 250.00 55.00

5 295.24 55.49 295.24 55.00

6 215.17 69.72 215.00 70.00

7 229.35 69.72 229.35 70.00

8 250.20 69.72 250.00 70.00

9 295.24 69.72 295.24 70.00

10 340.55 69.72 340.55 70.00

11 282.17 95.24 282.17 95.00

12 298.85 95.24 300.00 95.00

13 319.70 95.24 319.70 95.00

14 347.50 95.24 347.50 95.00

15 375.30 95.24 375.00 95.00

16 407.83 95.24 407.50 95.00

17 375.30 127.49 375.00 127.50

18 389.20 127.49 389.20 127.50

19 407.83 127.49 407.50 127.50

20 430.90 127.49 430.90 127.50

21 407.83 148.01 407.50 147.50

22 423.95 148.01 423.95 147.50

23 440.35 148.01 440.35 147.50

24 458.70 148.01 458.70 147.50

25 436.46 172.64 436.46 172.50

26 448.97 172.64 450.00 172.50

27 461.48 172.64 462.50 172.50

28 479.55 172.64 480.00 172.50

29 407.83 111.20 407.50 111.20

30 407.83 76.45 407.50 76.45

57
Table 10. Actual and Nominal Pressure Port Locations at Airplane Scale for 0.04-Scaled Model

Actual Nominal

Port FS BL FS BL

10 160.00 30.00 160.00 30.00

11 160.00 30.00 160.00 30.00

12 160.00 33.25 160.00 33.25

14 185.00 55.00 185.00 55.00

15 185.00 50.00 185.00 50.00

16 185.00 37.50 185.00 37.50

17 185.00 42.50 185.00 40.00

18 185.00 50.00 185.00 50.00

19 185.00 55.00 185.00 55.00

21 202.50 55.00 202.50 55.00

22 215.00 70.00 215.00 70.00

23 215.00 62.50 215.00 62.50

24 215.00 55.00 215.00 55.00

25 215.00 42.50 215.00 40.00

26 215.00 50.00 215.00 50.00

27 215.00 55.00 215.00 55.00

28 215.00 62.50 215.00 62.50

29 215.00 70.00 215.00 70.00

30 230.00 70.00 230.00 70.00

31 257.50 85.00 257.50 85.00

32 257.50 80.00 257.50 80.00

33 257.50 70.00 257.50 70.00

34 257.50 55.00 257.50 55.00

35 257.50 55.00 257.50 55.00

36 257.50 62.50 257.50 62.50

37 257.50 70.00 257.50 70.00

38 257.50 80.00 257.50 80.00

39 257.50 85.00 257.50 85.00

40 282.50 95.00 282.50 95.00

41 300.00 100.00 300.00 100.00

42 300.00 95.00 300.00 95.00

43 300.00 80.00 300.00 80.00

44 300.00 55.00 300.00 55.00

45 300.00 55.00 300.00 55.00

46 300.00 70.00 300.00 70.00

47 300.00 80.00 300.00 80.00

48 300.00 87.50 300.00 87.50

49 300.00 95.00 300.00 95.00

50 300.00 100.00 300.00 100.00

51 337.50 62.50 337.50 62.50

52 337.50 80.00 337.50 80.00

53 337.50 95.00 337.50 95.00

54 337.50 105.00 337.50 105.00

55 337.50 115.00 337.50 115.00

56 375.00 70.00 375.00 70.00

57 375.00 95.00 375.00 95.00

58 375.00 105.00 375.00 105.00

59 375.00 120.00 375.00 120.00

60 375.00 127.50 375.00 127.50

58
Table 10. Concluded

Actual Nominal

Port FS BL FS BL

61 390.00 127.50 390.00 127.50

62 407.50 147.50 407.50 147.50

63 407.50 137.50 407.50 137.50

64 407.50 105.00 407.50 105.00

65 407.50 55.00 407.50 55.00

67 407.50 95.00 407.50 95.00

68 407.50 120.00 407.50 120.00

69 407.50 127.50 407.50 127.50

70 407.50 137.50 407.50 137.50

71 407.50 147.50 407.50 147.50

72 425.00 62.50 425.00 62.50

73 425.00 147.50 425.00 147.50

74 425.00 160.00 425.00 160.00

75 437.50 147.50 437.50 147.50

76 437.50 160.00 437.50 160.00

77 437.50 172.50 437.50 172.50

78 450.00 80.00 450.00 80.00

79 450.00 147.50 450.00 147.50

80 450.00 160.00 450.00 160.00

81 450.00 172.50 450.00 172.50

82 462.50 172.50 462.50 172.50

83 462.50 147.50 462.50 147.50

84 462.50 105.00 462.50 105.00

86 462.50 105.00 462.50 105.00

88 462.50 172.50 462.50 172.50

89 492.50 172.50 492.50 172.50

90 492.50 147.50 492.50 147.50

91 492.50 105.00 492.50 105.00

92 492.50 55.00 492.50 55.00

93 492.50 80.00 492.50 80.00

94 492.50 120.00 492.50 120.00

95 492.50 147.50 492.50 147.50

96 492.50 160.00 492.50 160.00

59
60
Table 12. Boundary-Layer Velocity Proles

(a) Flight 135

Boundary-layer prole for run|

y, in. 6b at| 6b at| 10b at| 10b at| 11b at| 11b at| 12b at| 12b at|

RK 3 RK 5 RK 3 RK 5 RK 3 RK 5 RK 3 RK 5

0.02 0.690 0.718 0.692 0.711 0.709 0.711 0.709 0.702

0.06 0.742 0.781 0.738 0.789 7.750 0.781 0.753 0.779

0.10 0.760 0.821 0.764 0.828 0.773 0.819 0.782 0.815

0.14 0.781 0.846 0.780 0.850 0.792 0.838 0.799 0.839

0.18 0.798 0.874 0.799 0.875 0.808 0.862 0.813 0.862

0.28 0.848 0.909 0.848 0.907 0.951 0.989 0.858 0.895

0.38 0.874 0.932 0.878 0.926 0.888 0.915 0.892 0.915

0.48 0.898 0.940 0.910 0.939 0.913 0.922 0.939 0.929

0.58 0.928 0.951 0.938 0.948 0.942 0.938 0.951 0.935

0.68 0.942 0.952 0.957 0.953 0.966 0.942 0.971 0.941

0.88 0.970 0.930 0.980 0.960 0.980 0.954 0.990 0.951

1.08 0.982 0.963 0.988 0.968 0.998 0.960 0.999 0.960

1.28 0.991 0.972 0.993 0.974 0.995 0.968 0.988 0.972

1.48 0.995 0.980 0.999 0.982 0.999 0.972 0.995 0.978

1.88 1.000 1.000 1.000 1.000 1.000 1.000 1.000 1.000

(b) Flight 137

Boundary-layer prole for run|

y, in. 4b at| 4b at| 9b at| 9b at| 14b at| 14b at| 19b at| 19b at|

RK 3 RK 7 RK 3 RK 7 RK 3 RK 7 RK 3 RK 7

0.02 0.702 0.670 0.700 0.670 0.715 0.666 0.709 0.671

0.06 0.743 0.754 0.739 0.755 0.752 0.750 0.749 0.762

0.10 0.768 0.791 0.762 0.785 0.776 0.789 0.774 0.791

0.14 0.786 0.817 0.782 0.826 0.795 0.821 0.789 0.815

0.18 0.801 0.850 0.798 0.843 0.811 0.851 0.809 0.841

0.28 0.843 0.879 0.840 0.872 0.856 0.884 0.846 0.876

0.38 0.872 0.917 0.877 0.909 0.892 0.924 0.887 0.906

0.48 0.903 0.925 0.906 0.916 0.921 0.928 0.913 0.914

0.58 0.924 0.938 0.940 0.930 0.949 0.938 0.939 0.925

0.68 0.939 0.942 0.950 0.938 0.968 0.940 0.956 0.932

0.88 0.971 0.953 0.976 0.945 0.990 0.949 0.980 0.939

1.08 0.980 0.969 0.985 0.962 0.995 0.960 0.988 0.958

1.28 0.991 0.974 0.994 0.968 0.999 0.970 0.990 0.961

1.48 0.995 0.980 0.999 0.981 1.000 0.980 0.992 0.978

1.88 1.000 1.000 1.000 1.000 1.000 1.000 1.000

61
Table 12. Concluded

(c) Flight 138

Boundary layer prole for run|

y, in. 2c at| 2c at| 4b at| 4b at| 9b at| 9b at| 14b at| 14b at| 19b at| 19b at|

RK 3 RK 4 RK 3 RK 4 RK 3 RK 4 RK 3 RK 4 RK 3 RK 4

0.02 0.738 0.921 0.701 0.799 0.709 0.798 0.711 0.810 0.706 0.805

0.06 0.780 1.018 0.750 0.886 0.760 0.888 0.760 0.900 0.758 0.895

0.10 0.810 1.068 0.775 0.937 0.780 0.938 0.785 0.944 0.781 0.942

0.14 0.822 1.125 0.789 0.968 0.797 0.969 0.804 0.981 0.797 0.975

0.18 0.843 1.145 0.809 0.988 0.820 0.991 0.818 1.000 0.811 0.990

0.28 0.872 1.162 0.850 1.009 0.855 1.009 0.864 1.018 0.860 1.011

0.38 0.900 1.818 0.882 1.024 0.892 1.023 0.898 1.027 0.891 1.022

0.48 0.920 0.880 0.911 0.849 0.927 0.840 0.931 0.838 0.921 0.838

0.58 0.935 1.172 0.939 1.022 0.950 1.023 0.956 1.027 0.949 1.021

0.68 0.941 1.161 0.951 1.020 0.965 1.022 0.970 1.025 0.964 1.023

0.88 0.960 1.140 0.972 1.019 0.991 1.020 0.995 1.041 0.989 1.020

1.08 0.974 1.122 0.990 1.017 0.998 1.018 1.000 1.029 0.999 1.012

1.48 0.991 1.071 0.977 1.013 1.002 1.010 1.004 1.011 0.998 1.010

1.88 1.000 1.000 1.000 1.000 1.000 1.000 1.000 1.000 1.000 1.000

62
(a) Initial 
ow-visualization paint scheme on airplane near Langley Research Center.

(b) Modied 
ow -visualization paint scheme and video targets on airplane at Dryden Flight Research Center.

Figure 1. F-16XL-1 airplane in-
ight.

63
(c) Missiles, tufts, modied 
ow-visualization paint scheme, and video targets on airplane at Dryden Flight

Research Center.

Figure 1. Concluded.

64
32.4

(388.84)

FS

54.1

(649.86)

BL 137.096

17.6

(211.27)

5.40

(64.76)

Actuator

pod

Air dam

50°

70°

WL

BL

Figure 2. Three-view drawing of F-16XL-1 airplane. Linear dimensions are in feet (inches).

65
Figure 3. Artist's rendition of a High -Speed Civil Transport concept.

66
46.25 in

46.25 in

FS 0.000

FS 0.000

FS 215

52.8

FS 300

FS 375

FS 469

BL 41.5

BL 95

1 BL 172.5

Spacing, in.

3

12

Single point

Figure 4. Layout of photogrammetric targets on F-16XL-1 airplane. Approximately 32 additional single points

from previous measurements; approximately 30 additional single points of white 
ow viz targets.

67
Targets over wing

attachment bolts

Numerical surface

description

Typical line of

targets

(a) General.

Figure 5. Geometrical superposition of upper surface photogrammetric targets (depicted by lines) with

numerical surface description of F-16XL-1 airplane.

68
Numerical description

Targets

95

94

93

92

91

90

87

88

89

90 92 94 96 98 100 102

150

WL, in.

WL, in. 100

50

–150 0 50 100 150

BL, in.

BL, in.

–100 –50

(b) Cross section of airplane at FS 295 an d details on left wing; maximum dierence 0.09 in.

Figure 5. Continued.

69
200

150

100

50

110

105

100

95

90

150 200 250 300 350 400 450 500

Numerical description

Targets

165 170 175 180 185 190 195 200

WL, in.

WL, in.

FS, in.

FS, in.

(c) Cross section of port wing at BL 40 an d leading-edge details; maximum dierence 0.24 in.

Figure 5. Continued.

70
120

110

100

90

80

420

92

93

94

95

96

97

98

464 466 468 470 472 474 476

430 440 450 460 470 480 490 500 410 520

Numerical description

Targets

WL, in.

WL, in.

FS, in.

FS, in.

(d) Cross section of right wing at BL 165 and midchord details; maximum dierence 0.19 in.

Figure 5. Continued.

71
Numerical description

Line of targets

1.65 in.

(e) Cross section of forebody at BL 0 with details at nose and canopy.

Figure 5. Concluded.

72
37.19 in 37.19 in

37.19 in 37.19 in

Flush and belt static ports

Modified 

Hot-film sensors

Boundary-layer rakes

Leading

edge

Figure 6. Schematic of instrumentation on surface of airplane.

73
LE

LE flap

TE controls

Figure 7. Schematic of requested and actual pressure port layout and flight video targets on airplane.

Video targets

Requested pressure portsActual pressure ports

74
ESP (transducers)

Flush static ports with numberBelt static ports with number

21

21

33

BL 40.000

21

23 BL 36.0–39.0

9

Boundary-layer rakes (2)

(16 ports each)

Preston tubes (16)

(2 ports each)

Air dam

15

33

31

40

26

22

14

12

16

Figure 8. Details of requested complete pressure instrumentation suite and layout on airplane.

75
Figure 9. Front view of pressure instrumentation layout on right wing of F-16XL-1 airplane.

76
20

19

18

14

13

12

11

10

9

6

3

23

22

21

17

16

15

7

4

1

8

5

2

2.00

.25

.12 .36

.12

1.00

3.00

2.00

.06-in. connecting end tubes

.04-in. total-pressure tubes

.06-in. static-pressure tube

.04-in.

plate

Flatten to tube width/tube height = 3.00

1

2

3

4

5

6

7

8

9

10

11

12

*

*

*

*

*

*

*Port not used

*

0.020

0.033

0.047

0.060

0.073

0.087

0.100

0.113

0.127

0.140

0.180

0.280

13

14

15

16

17

18

19

20

21

22

23

0.380

0.480

0.580

0.680

0.780

0.880

1.080

1.280

1.480

1.680

1.880

Port Design

height

Port Design

height

Figure 10. Drawing of boundary-layer rake.

77
3

4

7

5

Boundary-layer rake

Modified Preston tubes

FS 330

Extension of hot-film gauge

line on upper surface

LE spar

FS 292.17, BL –100.51

LE

LE flap

(a) General arran gement.

FS 302.17, BL –52.93

(b) Rake 3.

Figure 11. Suite of instruments on F-16XL-1 left wing, including rake locations and predicted local 
ow near

surface.  = 13; M1 = 0:28; Rn = 46:1 106 .

78
FS 293.45, BL –76.22

(c) Rake 4.

FS 295.52, BL –94.33

(d) Rake 7.

Figure 11. Continued.

79
FS 294.59, BL –96.06

(e) Rake 5.

Figure 11. Concluded.

80
37.19 in 37.19 in

37.19 in 37.19 in

Camera 5

Camera 3

Camera 7

(HUD)

Camera 4

Camera 6

Figure 12. Locations of components of airborne video system installed on F-16XL-1 airplane.

81
Hot film

(a) General.

(b) Surface details.

Figure 13. Hot-lm belt layout on airplane.

82
(a) Initial.

(b) Current.

Figure 14. Grid on wing-fuselage upper surface.

83
(a) Initial.

(b) Current.

Figure 15. Grid on wing-fuselage lower surface.

84
Fuselage

Inboard upper wing

Outboard upper wing

Outboard leading edge

Inboard leading edge

Nose

Wake

(a) Inner.

Figure 16. CFD block structure layout for F-16XL-1 airplane.

85
(b) Outer.

Figure 16. Concluded.

86
0 2000 3000 4000 5000 6000 1000

–3.5

–3.0

–2.5

–2.0

–1.5

–1.0

–.5

0

0

.5

1.0

1.5

2.0

2.5

Cycles

CL, CD

CL

 CD

L2 (residual)

L2

(a) Rn = 38:97 106.

Figure 17. Reynolds number eect on convergence histories for two solutions.

87
0 2 000 4 000 6 000 8 000 10 000 12 000

–3.5

–4.0

–3.0

–2.5

–2.0

–1.5

–1.0

–.5

0

.5

–.5

1.0

1.5

Cycles

CL

CL

L2 (residual)

L2

(b) Rn = 1:12 106.

Figure 17. Concluded.

88
–Cp

Flt152.Run12b,up 

CFL3D

0 .2 .4 .6 .8 1.0

x/c

–1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

(a) BL 55.

 Flt152.Run12b,up 

 Flt152.Run12b,low 

 CFL3D 

0 .2 .4 .6 .8 1.0 –1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

–Cp

x/c

(b) BL 70.

Figure 18. Predicted and measured 
ight Cp distribution on F-16XL-1 airplane at FC 1 ( = 5:5; M1 = 0:52;

Rn = 77:71 106 ).

89
 Flt152.Run12b,up 

 Flt152.Run12b,low 

 CFL3D 

0 .2 .4 .6 .8 1.0 –1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

–Cp

x/c

(c) BL 80.

 Flt152.Run12b,up 

 Flt152.Run12b,low 

 CFL3D 

0 .2 .4 .6 .8 1.0 –1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

–Cp

x/c

(d) BL 95.

Figure 18. Continued.

90
 Flt152.Run12b,up 

 Flt152.Run12b,low 

 CFL3D 

0 .2 .4 .6 .8 1.0 –1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

–Cp

x/c

(e) BL 105.

 Flt152.Run12b,up 

CFL3D

0 .2 .4 .6 .8 1.0 –1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

–Cp

x/c

(f) BL 153.5.

Figure 18. Continued.

91
 Flt152.Run12b,up 

CFL3D

0 .2 .4 .6 .8 1.0 –1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

–Cp

x/c

(g) BL 184.5.

 Flt152.Run12b,up 

CFL3D

–Cp

0 .2 .4 .6 .8 1.0

2y/b

–1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

(h) FS 185.

Figure 18. Continued.

92
 Flt152.Run12b,up 

CFL3D

0 .2 .4 .6 .8 1.0 –1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

–Cp

2y/b

(i) FS 300.

 Flt152.Run12b,up 

CFL3D

0 .2 .4 .6 .8 1.0 –1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

–Cp

2y/b

(j) FS 337.5.

Figure 18. Continued.

93
 Flt152.Run12b,up 

CFL3D

0 .2 .4 .6 .8 1.0 –1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

–Cp

2y/b

(k) FS 375.

 Flt152.Run12b,up 

CFL3D

0 .2 .4 .6 .8 1.0

–1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

–Cp

2y/b

(l) FS 407.5.

Figure 18. Continued.

94
 Flt152.Run12b,up 

CFL3D

0 .2 .4 .6 .8 1.0

–1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

–Cp

2y/b

(m) FS 450.

 Flt152.Run12b,up 

CFL3D

0 .2 .4 .6 .8 1.0

–1.00

–.75

–.50

–.25

0

.25

.50

.75

1.00

1.25

–Cp

2y/b

(n) FS 492.5.

Figure 18. Continued.

95
Flight port location

Flight Cp value

1.20

–0.50–2.20

C

p

CFD

FC 1

(o) Upper surface Cp distribution.

Figure 18. Concluded.

96
 Flt144.Run3b,up 

CFL3D

0 .2 .4 .6 .8 1.0

–1.0

0

–.5

.5

1.0

1.5

2.0

–Cp

x/c

(a) BL 55.

 Flt144.Run3b,up 

 Flt144.Run3b,low 

CFL3D

0 .2 .4 .6 .8 1.0

–1.0

0

–.5

.5

1.0

1.5

2.0

–Cp

x/c

(b) BL 70.

Figure 19. Predicted and measured 
ight Cp distribution on F-16XL-1 airplane at FC 46 ( = 10:4;

M1 = 0:53; Rn = 46:90 106).

97
 Flt144.Run3b,up 

 Flt144.Run3b,low 

CFL3D

0 .2 .4 .6 .8 1.0

–1.0

0

–.5

.5

1.0

1.5

2.0

–Cp

x/c

(c) BL 80.

–1.0

0

–.5

.5

1.0

1.5

2.0

 Flt144.Run3b,up 

 Flt144.Run3b,low 

CFL3D

0 .2 .4 .6 .8 1.0

–Cp

x/c

(d) BL 95.

Figure 19. Continued.

98
–1.0

0

–.5

.5

1.0

1.5

2.0

 Flt144.Run3b,up 

 Flt144.Run3b,low 

CFL3D

0 .2 .4 .6 .8 1.0

–Cp

x/c

(e) BL 105.

 Flt144.Run3b,up 

CFL3D

0 .2 .4 .6 .8 1.0

–1.0

0

–.5

.5

1.0

1.5

2.0

–Cp

x/c

(f) BL 153.5.

Figure 19. Continued.

99
 Flt144.Run3b,up 

CFL3D

0 .2 .4 .6 .8 1.0

–1.0

0

–.5

.5

1.0

1.5

2.0

–Cp

x/c

(g) BL 184.5.

 Flt144.Run3b,up 

CFL3D

0 .2 .4 .6 .8 1.0

–1.0

0

–.5

.5

1.0

1.5

2.0

–Cp

2y/b

(h) FS 185.

Figure 19. Continued.

100
 Flt144.Run3b,up 

CFL3D

0 .2 .4 .6 .8 1.0

–1.0

0

–.5

.5

1.0

1.5

2.0

–Cp

2y/b

(i) FS 300.

 Flt144.Run3b,up 

CFL3D

0 .2 .4 .6 .8 1.0

–1.0

0

–.5

.5

1.0

1.5

2.0

–Cp

2y/b

(j) FS 337.5.

Figure 19. Continued.

101
Flt144.Run3b,up 

CFL3D

0 .2 .4 .6 .8 1.0

–1.0

0

–.5

.5

1.0

1.5

2.0

–Cp

2y/b

(k) FS 375.

 Flt144.Run3b,up 

CFL3D

0 .2 .4 .6 .8 1.0

–1.0

0

–.5

.5

1.0

1.5

2.0

–Cp

2y/b

(l) FS 407.5.

Figure 19. Continued.

102
 Flt144.Run3b,up 

CFL3D

0 .2 .4 .6 .8 1.0

–1.0

0

–.5

.5

1.0

1.5

2.0

–Cp

2y/b

(m) FS 450.

 Flt144.Run3b,up 

CFL3D

0 .2 .4 .6 .8 1.0

–1.0

0

–.5

.5

1.0

1.5

2.0

–Cp

2y/b

(n) FS 492.5.

Figure 19. Continued.

103
Flight port location

Flight Cp value

1.20

–1.00–3.20

C

p

FC 46

CFD

(o) Upper surface Cp distribution.

Figure 19. Concluded.

104
Flt144.Run9b,up_FC34

CFL3D_FC34

Flt144.Run4b,up_FC49

CFL3D_FC49

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

–Cp

x/c

(a) BL 55.

Flt144.Run9b,up_FC34

Flt144.Run9b,low_FC34 

CFL3D_FC34

Flt144.Run4b,up_FC49

Flt144.Run4b,low_FC49

CFL3D_FC49

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

–Cp

x/c

(b) BL 70.

Figure 20. Predicted and measured 
ight Cp distribution on F-16XL-1 airplane at nom = 13 for FC 34

( = 13:5; M1 = 0:37; Rn = 40:05 106 ) and FC 49 ( = 13:0; M1 = 0:44; Rn = 38:97 106).

105
Flt144.Run9b,up_FC34

Flt144.Run9b,low_FC34 

CFL3D_FC34

Flt144.Run4b,up_FC49

Flt144.Run4b,low_FC49

CFL3D_FC49

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

–Cp

x/c

(c) BL 80.

Flt144.Run9b,up_FC34

Flt144.Run9b,low_FC34 

CFL3D_FC34

Flt144.Run4b,up_FC49

Flt144.Run4b,low_FC49

CFL3D_FC49

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

–Cp

x/c

(d) BL 95.

Figure 20. Continued.

106
Flt144.Run9b,up_FC34

Flt144.Run9b,low_FC34 

CFL3D_FC34

Flt144.Run4b,up_FC49

Flt144.Run4b,low_FC49

CFL3D_FC49

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

–Cp

x/c

(e) BL 105.

Flt144.Run9b,up_FC34

CFL3D_FC34

Flt144.Run4b,up_FC49

CFL3D_FC49

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

–Cp

x/c

(f) BL 153.5.

Figure 20. Continued.

107
Flt144.Run9b,up_FC34

CFL3D_FC34

Flt144.Run4b,up_FC49

CFL3D_FC49

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

–Cp

x/c

(g) BL 184.5.

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

Flt144.Run9b,up_FC34

CFL3D_FC34

Flt144.Run4b,up_FC49

CFL3D_FC49

–Cp

2y/b

(h) FS 185.

Figure 20. Continued.

108
Flt144.Run9b,up_FC34

CFL3D_FC34

Flt144.Run4b,up_FC49

CFL3D_FC49

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

–Cp

2y/b

(i) FS 300.

Flt144.Run9b,up_FC34

CFL3D_FC34

Flt144.Run4b,up_FC49

CFL3D_FC49

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

–Cp

2y/b

(j) FS 337.5.

Figure 20. Continued.

109
Flt144.Run9b,up_FC34

CFL3D_FC34

Flt144.Run4b,up_FC49

CFL3D_FC49

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

–Cp

2y/b

(k) FS 375.

Flt144.Run9b,up_FC34

CFL3D_FC34

Flt144.Run4b,up_FC49

CFL3D_FC49

–1.0

0 .2 .4 .6 .8 1.0

–.5

0

.5

1.0

1.5

2.0

2.5

–Cp

2y/b

(l) FS 407.5.

Figure 20. Continued.

110
Flt144.Run9b,up_FC34

CFL3D_FC34

Flt144.Run4b,up_FC49

CFL3D_FC49

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

–Cp

2y/b

(m) FS 450.

Flt144.Run9b,up_FC34

CFL3D_FC34

Flt144.Run4b,up_FC49

CFL3D_FC49

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

–Cp

2y/b

(n) FS 492.5.

Figure 20. Con cluded.

111
 Flt144.Run16b,up 

 BART,up 

 BART,low 

CFL3D_FC25

CFL3D_WT-BART

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

–Cp

x/c

(a) BL 55.

 Flt144.Run16b,up 

 Flt144.Run16b,low 

 BART,up 

 BART,low 

 CFL3D_FC25

 CFL3D_WT-BART

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

–Cp

x/c

(b) BL 70.

Figure 21. Predicted and measured 
ight and WT Cp distributions on F-16XL -1 airplane at nom = 20 for

FC 25 ( = 19:8; M1 = 0:24; Rn = 32:22 106), BART ( = 20:0; M1 = 0:15; Rn = 1:06 106), and

CFD for BART ( = 20:0; M1 = 0:10; Rn = 1:12 106).

112
 Flt144.Run16b,up 

 Flt144.Run16b,low 

 BART,up 

 BART,low 

 CFL3D_FC25

 CFL3D_WT-BART

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

–Cp

x/c

(c) BL 80.

 Flt144.Run16b,up 

 Flt144.Run16b,low 

 BART,up 

CFL3D_FC25

CFL3D_WT-BART

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

–Cp

x/c

(d) BL 95.

Figure 21. Continued.

113
 Flt144.Run16b,up 

 Flt144.Run16b,low 

 BART,up 

 BART,low 

 CFL3D_FC25

 CFL3D_WT-BART

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

–Cp

x/c

(e) BL 105.

 Flt144.Run16b,up 

CFL3D_FC25

CFL3D_WT-BART 

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

–Cp

x/c

(f) BL 153.5.

Figure 21. Continued.

114
 Flt144.Run16b,up 

CFL3D_FC25

CFL3D_WT-BART 

0 .2 .4 .6 .8 1.0 –1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

–Cp

x/c

(g) BL 184.5.

 Flt144.Run16b,up 

 BART,up 

 BART,low 

CFL3D_FC25

CFL3D_WT-BART 

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

–Cp

2y/b

(h) FS 185.

Figure 21. Continued.

115
 Flt144.Run16b,up 

 BART,up 

 BART,low 

CFL3D_FC25

CFL3D_WT-BART

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

–Cp

2y/b

(i) FS 300.

 Flt144.Run16b,up 

 BART,up 

CFL3D_FC25

CFL3D_WT-BART

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

–Cp

2y/b

(j) FS 337.5.

Figure 21. Continued.

116
 Flt144.Run16b,up 

 BART,up 

CFL3D_FC25

CFL3D_WT-BART

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

–Cp

2y/b

(k) FS 375.

 Flt144.Run16b,up 

 BART,up 

 BART,low 

CFL3D_FC25

CFL3D_WT-BART

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

–Cp

2y/b

(l) FS 407.5.

Figure 21. Continued.

117
 Flt144.Run16b,up 

 BART,up 

CFL3D_FC25

CFL3D_WT-BART

0 .2 .4 .6 .8 1.0

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

–Cp

2y/b

(m) FS 450.

–1.0

–.5

0

.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

 Flt144.Run16b,up 

 BART,up 

 BART,low 

CFL3D_FC25

CFL3D_WT-BART

0 .2 .4 .6 .8 1.0

–Cp

2y/b

(n) FS 492.5.

Figure 21. Continued.

118
BART; α = 20°; M∞ < 0.20

1.00

–3.50–8.00

Cp

CFD

PSP

(o) Upper surface Cp distribution of BART PSP results and CFD.

Figure 21. Concluded.

119
 Flt152.Run5b,up.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

x/c

(a) BL 40.

 Flt152.Run5b,up.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

x/c

(b) BL 55.

Figure 22. Predicted and measured 
ight Cp distribution on F-16XL-1 airplane for FC 70 ( = 4:4

; M1 = 0:97;

Rn = 88:77 106 ).

120
 Flt152.Run5b,up.0.97 

 Flt152.Run5b,low.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

x/c

(c) BL 70.

 Flt152.Run5b,up.0.97 

 Flt152.Run5b,low.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

x/c

(d) BL 80.

Figure 22. Continued.

121
 Flt152.Run5b,up.0.97 

 Flt152.Run5b,low.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

x/c

(e) BL 95.

 Flt152.Run5b,up.0.97 

 Flt152.Run5b,low.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

x/c

(f) BL 105.

Figure 22. Continued.

122
 Flt152.Run5b,up.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

x/c

(g) BL 127.5.

 Flt152.Run5b,up.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

x/c

(h) BL 153.5.

Figure 22. Continued.

123
 Flt152.Run5b,up.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

x/c

(i) BL 184.5.

 Flt152.Run5b,up.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

2y/b

(j) FS 185.

Figure 22. Continued.

124
 Flt152.Run5b,up.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

2y/b

(k) FS 300.

 Flt152.Run5b,up.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

2y/b

(l) FS 337.5.

Figure 22. Continued.

125
 Flt152.Run5b,up.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

2y/b

(m) FS 375.

 Flt152.Run5b,up.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

2y/b

(n) FS 407.5.

Figure 22. Continued.

126
 Flt152.Run5b,up.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–.75

–.50

–.25

0

.25

.50

.75

–Cp

2y/b

(o) FS 450.

–.75

–.50

–.25

0

.25

.50

.75

 Flt152.Run5b,up.0.97 

CFL3D

0 .2 .4 .6 .8 1.0

–Cp

2y/b

(p) FS 462.5.

Figure 22. Continued.

127
 Flt152.Run5b,up.0.97 

CFL3D

0 .2 .4 .6 .8 1.0 –.75

–.50

–.25

0

.25

.50

.75

–Cp

2y/b

(q) FS 492.5.

Figure 22. Continued.

128
Flight port location

Flight Cp value

1.200.00

–1.20

Cp

CFD

FC 70

(r) Upper surface Cp distribution.

Figure 22. Concluded.

129
Flight

CFD

FC 70

1.20

Cp

0.00

–1.20

Flight port location

Figure 23. Cp,u contours on F-16XL-1 airplane for FC 70 (α = 4.4°; M∞ = 0.97; Rn = 88.77 × 1

130
.2

.4

.6

.8

1.0

1.2

V/VRE

0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0

y, in.

 Flt145.Run16b Prediction 

 Flt135.Run12b Measured 

 Flt137.Run19b Measured 

 Flt138.Run19b Measured 

Rake 3

LE

(a) Rake 3; FS 302.17, BL 52.93.

LE

 Flt145.Run16b Prediction 

 Flt138.Run19b Measured 

.2

.4

.6

.8

1.0

1.2

V/VRE

0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0

y, in.

Rake 4

(b) Rake 4; FS 293.45, BL 76.22.

Figure 24. Predicted and measured velocity proles for boundary-layer rakes on F-16XL-1 airplane for FC 7

(nom = 13; nom = 0; M1 = 0:29; h = 5000 ft; Rn = 44:4 106).

131
LE

 Flt145.Run16b Prediction 

 Flt137.Run19b Measured 

.2

.4

.6

.8

1.0

1.2

V/VRE

0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0

y, in.

Rake 7

(c) Rake 7; FS 295.52, BL 94.33.

LE

 Flt145.Run16b Prediction 

 Flt135.Run12b Measured 

.2

.4

.6

.8

1.0

1.2

V/VRE

0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0

y, in.

Rake 5

(d) Rake 5; FS 294.59, BL 96.06.

Figure 24. Concluded.

132
LE

 Flt137.Run19b, Rake 7

 Flt135.Run12b, Rake 5

.2

.4

.6

.8

1.0

1.2

V/VRE

0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0

y, in.

7 5

Rakes

(a) Measured velocity proles.

Rake 5 

Rake 7 

Vortex locations using

stagnation pressure

contours and streamline

ribbons

(b) Some predicted global features at FS 295.

Figure 25. O-surface 
ow features for bou ndary-layer rake locations 7 and 5 on F-16XL-1 airplane for FC 7

(nom = 13; nom = 0; M1 = 0:29; h = 5000 ft; Rn = 44:4 106; FS  295).

133
Third vortex 

Secondary vortex

Primary vortex

(c) Origination of predicted vortex systems inboard of wing cran k.

Figure 25. Concluded.

134
–120 –100 –80 –60 –40

BL, in.

0

.002

.004

.006

.008

.010

.012

.014

cf

Flt131.Run14b,α11.85°,M0.36, R46.8, EFD_FC19

α11.89°,M0.30, R44.4, CFD_FC7

Modified Preston

tubes

LE

Figure 26. Predicted and measured cf on F-16XL -1 airplane at FC 330 for M1;avg = 0:33, avg = 11:9

,

nom = 0, Rn;avg = 45:6 106.

135
(a) Tuft images pro jected from three cameras onto aircraft grid.

Figure 27. Flight tuft data from three cameras on F-16XL-1 airplane and CFD solution at FC 46 ( = 10:4;

M1 = 0:53; Rn = 46:90 106).

136
(b) Combination of tuft images and CFD surface streamlines.

Figure 27. Continued.

137
(c) Vortex systems determined from CFD stagnation pressures (PLOT3D) at 0.78.

Figure 27. Continued.

138
(d) Combination of tuft images, streamlines, and vortex systems.

Figure 27. Concluded.

139
–10 –5 0 5 10 15 20 25 30 35

α, deg

–.4

0

.4

.8

1.2

1.6

 CL

 CL, Cm

 Cm

(a) CD and Cm versus .

–.4 0 .4 .8 1.2 1.6

CL

–.2

0

.2

.4

.6

.8

 CD

 Cm

 CL, Cm

(b) CD and Cm versus CL.

Figure 28. Longitudinal aerod ynamic characteristics of 0.18-scaled model of F-16XL-1 airplane with air dams

and missiles on and controls unde
ected at M1 = 0:07 and Rn = 2:1 106 .

140
 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

0 .2 .4 .6 .8 1.0

x/c

–Cp

–.5

0

.5

1.0

1.5

(a) BL 55.

0 .2 .4 .6 .8 1.0

 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

–.5

0

.5

1.0

1.5

x/c

–Cp

(b) BL 70.

Figure 29. Eect of Rn variation on local Cp for 0.04- (BART) and 0.18-scaled ( 30 by 60) models of F-16XL-1

airplane in BART and 30- by 60-Foot Tunnel, respectively, with air dams and missiles for  = 5 at

M1  0:2.

141
0 .2 .4 .6 .8 1.0

 BART,up,R0.5 

 BART,up,R1.0 

 BART,up,R1.12 

 30_by_60,up,R2.1 

–.5

0

.5

1.0

1.5

x/c

–Cp

(c) BL = 95.

0 .2 .4 .6 .8 1.0

 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

–.5

0

.5

1.0

1.5

x/c

–Cp

(d) BL = 147.5.

Figure 29. Concluded.

142
–.5

1.0

0

1.5

.5

2.0

0 .2 .4 .6 .8 1.0

 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

x/c

–Cp

(a) BL 55.

0 .2 .4 .6 .8 1.0

 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

–.5

1.0

0

1.5

.5

2.0

x/c

–Cp

(b) BL 70.

Figure 30. Eect of Rn variation on local Cp for 0.04- (BART) and 0.18-scaled ( 30 by 60) models of F-16XL-1

airplane in BART and 30- by 60-Foot Tunnel, respectively, with air dams and missiles for  = 13 at

M1  0:2.

143
0 .2 .4 .6 .8 1.0

 BART,up,R0.5 

 BART,up,R1.0 

 BART,up,R1.12 

 30_by_60,up,R2.1 

–.5

1.0

0

1.5

.5

2.0

x/c

–Cp

(c) BL 95.

0 .2 .4 .6 .8 1.0

 BART,up,R0.5 

 BART,up,R1.0 

 BART,up,R1.12 

 30_by_60,up,R2.1 

–.5

1.0

0

1.5

.5

2.0

x/c

–Cp

(d) BL 147.5.

Figure 30. Concluded.

144
0 .2 .4 .6 .8 1.0

–1

0

1

2

3

4

 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

x/c

–Cp

(a) BL 55.

0 .2 .4 .6 .8 1.0

 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

–1

0

1

2

3

4

x/c

–Cp

(b) BL 70.

Figure 31. Eect of Rn variation on local Cp for 0.04- (BART) and 0.18-scaled ( 30 by 60) models of F-16XL-1

airplane in BART and 30- by 60-Foot Tunnel, respectively, with air dams and missiles for  = 20 at

M1  0:2.

145
0 .2 .4 .6 .8 1.0

 BART,up,R0.5 

 BART,up,R1.0 

 BART,up,R1.12 

 30_by_60,up,R2.1 

–1

0

1

2

3

4

x/c

–Cp

(a) BL 95.

0 .2 .4 .6 .8 1.0

 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

–1

0

1

2

3

4

x/c

–Cp

(b) BL 147.5.

Figure 31. Concluded.

146
0 .2 .4 .6 .8 1.0

2y/b

–Cp

–.5

0

.5

1.0

1.5

 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

(a) FS 215.

0 .2 .4 .6 .8 1.0

 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

–.5

0

.5

1.0

1.5

2y/b

–Cp

(b) FS 407.5.

Figure 32. Eect of Rn variation on local Cp for 0.04- (BART) and 0.18-scaled ( 30 by 60) models of F-16XL-1

airplane in BART and 30- by 60-Foot Tunnel, respectively, with air dams and missiles for  = 5 at

M1  0:2.

147
0 .2 .4 .6 .8 1.0

–.5

0

.5

1.0

1.5

2.0

 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

2y/b

–Cp

(a) FS 215.

0 .2 .4 .6 .8 1.0

 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

–.5

0

.5

1.0

1.5

2.0

2y/b

–Cp

(b) FS 407.5.

Figure 33. Eect of Rn variation on local Cp for 0.04- (BART) and 0.18-scaled ( 30 by 60) models of F-16XL-1

airplane in BART and 30- by 60-Foot Tunnel, respectively, with air dams and missiles for  = 13 at

M1  0:2.

148
0 .2 .4 .6 .8 1.0

–1

0

1

2

3

4

 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

2y/b

–Cp

(a) FS 215.

0 .2 .4 .6 .8 1.0

 BART,up,R0.5 

 BART,low,R0.5 

 BART,up,R1.0 

 BART,low,R1.0 

 BART,up,R1.12 

 BART,low,R1.12 

 30_by_60,up,R2.1 

–1

0

1

2

3

4

2y/b

–Cp

(b) FS 407.5.

Figure 34. Eect of Rn variation on local Cp for 0.04- (BART) and 0.18-scaled ( 30 by 60) models of F-16XL-1

airplane in BART and 30- by 60-Foot Tunnel, respectively, with air dams and missiles for  = 20 at

M1  0:2.

149
α = –2.27° α = 2.04°

α = –0.14° α = 4.27°

Cp +

0

–

+

0

–

Cp

Shock location

Cp Cp

Shock location

+

0

–

+

0

–

(a) 2:27    4:27.

Figure 35. Eect of  on Cp;u contours for 0.11-scaled model in Ames 11-Foot Tunnel at M1 = 0:95 and

Rn = 2:3 106.

150
α = 6.54° α = 11.07°

α = 8.78° α = 11.35°

+

–

Cp +

0 0

–

Cp

+

–

Cp +

0 0

–

Cp

Shock location

Shock location

Shock location

Shock location

(b) 6:54    11:35.

Figure 35. Continued.

151
α = 20.04° α = 15.44°

α = 17.79°

+

–

Cp +

0 0

–

Cp

+

0

–

Cp

(c) 15:44    20:04.

Figure 35. Concluded.

152
Reattachment path

Core path

(a)  = 10; M1 = 0:15; Rn = 1 106.

Figure 36. Superp osition of projected core path and reattachment path onto oil 
ow image for F-16XL-1 BART

model without air dams.

153
Reattachment line

Reattachment path

Core path

(b)  = 13; M1 = 0:07; Rn = 0:5 106.

Figure 36. Continued.

154
Reattachment path

Core path

(c)  = 13; M1 = 0:15; Rn = 1 106.

Figure 36. Continued.

155
Reattachment path

Core path

(d)  = 15; M1 = 0:07; Rn = 0:5 106.

Figure 36. Continued.

156
Reattachment path

Core path

(e)  = 15; M1 = 0:15; Rn = 1 106.

Figure 36. Concluded.

157
Figure 37. Su perp osition of liquid crystal and tuft image data for F-16XL-1 airplane at   13, M1 = 0:28,

and Rn = 47 106.

158
Data

fusion

Figure 38. Su perp osition of liquid crystal and tuft image data for F-16XL-1 airplane at   13, M1 = 0:28,

and Rn = 47 106.

159
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REPORT DOCUMENTATION PAGE

February 2001 Technical Publication

Flight, Wind-Tunnel, and Computational Fluid Dynamics Comparison for

Cranked Arrow Wing (F-16XL-1) at Subsonic and Transonic Speeds WU 522-31-31-03

John E. Lamar, Clifford J. Obara, Bruce D. Fisher, and David F. Fisher

L-18018

NASA/TP-2001-210629

John E. Lamar, Clifford J. Obara, and Bruce D. Fisher: Langley Research Center, Hampton, VA; David F. Fisher:

Dryden Flight Research Center, Edwards, CA.

Geometrical, flight, computational fluid dynamics (CFD), and wind-tunnel studies for the F-16XL-1 airplane are

summarized over a wide range of test conditions. Details are as follows: (1) For geometry, the upper surface of the

airplane and the numerical surface description compare reasonably well. (2) For flight, CFD, and wind-tunnel surface pressures, the comparisons are generally good at low angles of attack at both subsonic and transonic speeds;

however, local differences are present. In addition, the shock location at transonic speeds from wind-tunnel pressure contours is near the aileron hinge line and generally is in correlative agreement with flight results. (3) For

boundary layers, flight profiles were predicted reasonably well for attached flow and underneath the primary vortex

but not for the secondary vortex. Flight data indicate the presence of an interaction of the secondary vortex system

and the boundary layer and the boundary-layer measurements show the secondary vortex located more outboard

than predicted. (4) Predicted and measured skin friction distributions showed qualitative agreement for a two vortex

system. (5) Web-based data-extraction and computational-graphical tools have proven useful in expediting the preceding comparisons. (6) Data fusion has produced insightful results for a variety of visualization-based data sets.

F-16XL-1; Flow physics; Subsonic; Transonic; Data fusion 168

A08

NASA Langley Research Center

Hampton, VA 23681-2199

National Aeronautics and Space Administration

Washington, DC 20546-0001

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