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AIAA-2003-0752
X-29 High Alpha Test in the National TransonicFacility (Invited)
Pamela J. Underwood, ViGYAN, Inc.
Hampton, VA 23666
Lewis R. Owens, Richard A. Wahls, and Susan Williams
NASA Langley Research CenterHampton, VA 23681
41st AIAA Aerospace Sciences Meeting & Exhibit6- 9 January 2003
Reno, NevadaI I
For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics1801 Alexander Bell Drive, Suite 500, Reston, VA 20191-4344
AIAA-2003-0752
X-29 High Alpha Test in the National Transonic Facility
Pamela J. Under wood*
ViGYAN, Inc., Hampten, VA 23666
Lewis R. Owens 1,Richard A. W"ahls t, Susan Williams _
NASA Langley Research CenteL Hampton, VA 23681
Abstract Introduction
This paper describes the X-29A research program atthe National Transonic Facility. This wind tunnel
test leveraged the X-29A high alpha flight test
program by enabling ground-to-flight correlationstudies with an emphasis on Reynolds number
effects. The background and objectives of this testprogram, as well as the comparison of high Reynolds
number wind tunnel data to X-29A flight test data arepresented. The effects of Reynolds number on the
forebody pressures at high angles of attack are alsopresented• The purpose of this paper is to document
this test and serve as a reference for future ground-to-flight correlation studies, and high angle-of-attack
investigations. Good ground-to-flight correlationswere observed for angles of attack up to 50 °, and
Reynolds number effects were also observed.
Nomenclature
c mean aerodynamic chord, innormal force coefficientC_
CrCv
DARPA
ESP
M
pressure coefficientside force coefficient
Defense Advanced Research
Projects Agencyelectronically scanned pressure
de_rees Fahrenheitfuselage or body length, in
Mach number
NTF National Transonic FacilityP!
Rc
ReD
USAF
total pressure, psiReynolds number based on mean
aerodynamic chordReynolds number based on
forebody diameterWing reference area, inz
total temperature, °FUnited States Air Force
axial distance from nose apex, in
angle of attack, de_circumferential angle, deg
* Research Engineer, Member AIAA
Aerospace Engineer, Flow Physics and Control Branch, SeniorMember AIAA
Asst. Head, Configuration Aerodynamics Branch, AssociateFellow AIAA
+ Aerospace Engineer
This material is declared work of the U.S. Government and is not
subject to copyright protection in the United States.
-_joint research program to investigate the high
ingle-of-attack performance potential of the X-29A:brward swept wing fighter commenced in the
1980"s. Primary partners in this joint program were_ASA, Grumman Aerospace, the United States Airf:orce (USAF), and the Defense Advanced Research
Projects Agency (DARPA). During the high angle-
of-attack flight test, data were obtained to supportaerodynamics, flow visualization, control systems,
imndling qualities, and maneuverability research. Amore extensive description of this flight program maybe found in reference 1. The National Transonic
Facility (NTF) X-29 High Alpha test was established
to augment this flight research program by enablingground-to-flight correlation studies with emphasis on
Reynolds number effects. Portions of the data from1his wind tunnel test have been previously published
in related amcles. +'- The purpose of this paper is tofurther document this test and serve as a reference for
future ground-to-flight correlation studies, and highangle-of-attack investigations.
Aircraft DescriptionThe X-29A is a single place research airplane with a29.27 ° forward leading edge sweep wing, close
coupled variable incidence canards, and full spandual hinged flaperons. 4 The Grumman Aerospace
designed X-29A is shown in figure 1. The forwardswept wing is an aeroelastically tailored composite
structure with a supercritical airfoil and a fixedleading edge. A 75-inch long nose boom, and 24-
inch long nose strakes are positioned at the noseapex. Side and bottom views of the nose apex are
shown in figure 2.
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American Institute of Aeronautics and Astronautics
Figure 1: Grumman Aerospace X-29A
Angie-of-altackvanes_T'.L.L"
9o° 27oo :
............... l °f .............!............i,,0o
e=oo e=o°x/I = 0201
x/I=0.136
Figure 3: X-29A Research Aircraft ForebodyPressure Locations
The distribution of these static pressure rows is also
shown in figure 3, where 0° is the windward side ofthe fuselage, 90 ° is the starboard side, and 180 ° is the
leeward side. In addition to these static pressureorifices on the forebody, the research airplane was
also instrumented with three angle of attack vanesand one angle of sideslip vane on the nose boom.
Facility DescriptionThe NTF is a unique transonic wind tunnel designedto conduct full-scale flight Reynolds number testing
through the use of high pressures and cryogenictemperatures. This is a fan driven, closed circuit
wind tunnel with an 8.2-foot by 8.2-foot and 25-footlong test section with a slotted ceiling and floor. A
planform view of the tunnel is provided in figure 4.
Figure 2: Side and Bottom View of Nose Apex
The nose boom tapered from a 0.88-inch diameter atthe tip to a 3.5-inch diameter at the nose apex, andthe nose strakes were 1.5 inches wide at the nose
apex and 2.5 inches wide at the downstream end.As shown in figure 3 the research airplane wasinstrumented with four circumferential rows of static
pressure orifices.
200 --- - *]
! Low-speed diffuser 19.7-dia fan ]Turn 3 i
, , ] Turn 2
_'(. ! • " n, nf t-t , "7 "_
48.6 ] _ _ " _95:lcontraction_ l_Jl
_: n f, a
Turn 4 _ Screens _ _ High-speed diffuser
_-Cooling coil / 27-dia plenum 2.6" halt-angle
± Wide-angle diffuser / Slotted test section8.2 by 8.2
Figure 4: Planform View of the NTF
The tunnel operates using either dry air or gaseousnitrogen as the test medium. During air operations
the tunnel pressure is used to control Reynolds
number, while in the cryogenic nitrogen mode thetunnel temperature and pressure are used to controlReynolds number. The NTF affords a test
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American Institute of Aeronautics and Astronautics
environment in which the Mach number and chord
Reynolds number are identical for the scale model in
the test section and the full-scale aircraft in flight.The NTF is capable of an absolute pressure range
from 15 psia to 125 psia, a temperature range from
-320°F to 150°F, a Mach number range from 0.2 to
1.2, and a maximum Reynolds number of 146x 106
per ft at Mach 1. Typical tests use a temperature
range from -250°F to 120°F. A more extensive
facility description can be found in reference 5.
NTF X-29 Test Program
The primary test objectives were to compare the NTFhigh Reynolds number forebody pressure data to the
data obtained during the X-29A high alpha flight test,and to assess the Reynolds number effects on the
forebody flow at high angles of attack. The effect offixing transition on the forebody was also studied
during this test program.
1/16 th Scale X-29 Model
The NTF X-29 model is a 1/16 th scale representation
of the research airplane. All of the components of theX-29A research airplane were accurately scaled for
the NTF model except for the thickness of the nosestrakes. At 1/16 th scale the model nose strakes should
have been 0.0075 inches thick, but were actually 0.03
inches thick due to NTF model strengthrequirements. Pertinent model geometry is given infigure 5.
['he 1/16 th scale NTF model featured flow through
_nlets positioned on either side of the forebody justOrward of the canards that combined to form a singleexhaust at the back of the model. A flow shield was
ncluded to isolate the balance from the interior duct!low in this model.
£he contour tolerance of the wing, canard, andvertical tail was _+0.002 inches. The fuselage
forebody tolerance was __.0.004 inches, while the
remaining fuselage tolerance was approximately_ 0.004 inches to - 0.006 inches. The model was
built of 18% nickel maraging steel (C type) with asurface finish of approximately 10 microinches
(RMS). The model was composed of separablecomponents to allow testing of multiple
configurations. The flaperons, aft body strakes,rudder, and canards were all designed to be set atdiscrete angles. The 1/16 th scale NTF X-29 model is
shown in figure 6 with all its control surface
components. During this NTF test only the canardangle was varied. The model canard was designed to
accommodate five discrete angle positions (-20 °,-25 ° , -30 °, -35 °, -60°), and was set to match the flight
test conditions as closely as possible.
Adjustable Control Surfaces
Flow Through _/ \\
Inlets _.___r_n l
_ --_ __::_ 20.4 in
c = 5.41 in _ [
S = 104.1 in z
36 in
Figure 5:1/16 th Scale X-29 Model Geometry
Figure 6: NTF X-29 Model with Control Surfaces
A unique high alpha sting was used with the X-29model. This sting was designed to accommodate anangle of attack range from -7 ° to 74 ° using three
primary knuckle positions. A sketch of this high
alpha sting showing the three possible knuckleD positions is given in Figure 7. As seen in this sketch,
only the second knuckle position keeps the modelcenter of rotation aligned with the tunnel centerline,
and positions #1 and #3 would present the modelpositions most susceptible for wall interference.Figure 8 shows the 1/16 th scale X-29 model mounted
on the high alpha sting in the NTF test section.
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American Institute of Aeronautics and Astronautics
m
/.//-/_',\ :osition #3
Position #1 / "-. -.
Figure 7: General Arrangement of NTF High
Alpha Sting Depicting Three Knuckle Positions
(Not to Scale)
Figure 8:X-29 Model Mounted in NTF Test Section
(Knuckle Position #3)
Instrumentation and Measurement Corrections
The NTF model was instrumented with one
circumferential row of static pressure orifices on the
forebody at station x/l=0.136. This forebody location
was chosen to correspond with one of the X-29A
research airplane static pressure row locations (see
figure 3). The circumferential distribution of the
model static pressure orifices was also positioned to
match the research airplane as closely as possible.
Table 1 gives the circumferential pressure orifice
locations for both the X-29 research airplane and the
X-29 NTF model at x/l=O. 136. Again, 0 ° is the
windward side of the fuselage, 90 ° is the starboard
side, and 180 ° is the leeward side.
electronically scanned pressure (ESP) module with a
full-scale pressure range of_+ 15 psid. The quoted
(worst case) accuracy of the ESP module was
approximately 0.20 percent of full scale or _+0.030
psi (Cp variation _-q).06 at the lowest dynamic
pressure condition). For reference, the X-29A high
alpha flight test forebody pressure data were obtained
using _+ 1.5 psi differential pressure transducers with
an estimated accuracy of approximately _+0.007 psi. 4
The model aerodynamic force and moment data were
obtained with a six component unheated strain gage
balance. The balance maximum load capacity and
quoted accuracies are given in Table 2. The axial
force and moment data acquired were inconsistent,
and deemed corrupt. These data show signs of
interference on the balance most likely due to the
tightly packed instrumentation within the model.
Normal and side force data were less sensitive to this
adverse effect, and are presented herein. The main
objective was to compare forebody pressures with
available flight data.
Table h Circumferential Static Pressure Locations,
x/1=0.136
NTF Flight0 (deg) 0 (deg)
24.2 n/a
47.6 n/a
59.6 60.0
n/a 66.0
72.0 72.0
77.9 78.0
83.9 n/a
89.8 90.0
95.0 95.0
99.9 100.0
103.7 105.0
107.7 108.0
114.1 111.1
119.9 120.0
125.9 126.0
n/a 129.0
131.9 132.0
n/a 135.0
138.0 138.0
n/a 141.0
143.9 144.0
n/a 147.0
150.0 150.0
n/a 153.0
155.9 156.0
162.1 165.0
167.9 168.0
n/a 171.0
174.0 174.0
180.0 n/a
The NTF model forebody pressure data were
obtained through the use of one internal 48 port
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American Institute of Aeronautics and Astronautics
NTF Flight
0 (deg) 0 (deg)
185.9 186.0
n/a 189.0
191.9 192.0
197.7 195.0
203.6 204.0
n/a 207.0
209.8 210.0
n/a 213.0
215.5 216.0
n/a 219.0
221.4 222.0
n/a 225.0
227.3 228.0
n/a 231.0
234.7 234.0
n/a 237.0
240.5 240.0
246.2 249.0
252.3 252.0
n/a 255.0
260.3 260.0
265.2 265.0
270.0 270.0
276.0 276.0
288.0 288.0
n/a 294.5
299.5 300.0
336.1 n/a
360.0 n/a
Table 2: NTF Balance Load Capacity
Maximum Full-ScaleComponent Load Accuracy
Normal Force _+2500 lb 0.10 %Axial Force -4-350 lb 0.26 %
_ 5000 in-lb 0,11%Pitching MomentRolling MomentYawin 8 Moment
Side Force
_+2500 in-lb 0.40 %_+4000 in-lb 0.18 %
+ 1000 lb 0.35 %
_hese figures were obtained without fixing transitionm the 1/16 thscale NTF model. Figures 9a and 9b.tre for test conditions of 0.25 Mach number (M),
Reynolds number based on mean aerodynamic chord
Pc) of approximately 6.6 million, and angles of
attack (c_) of approximately 30 and 35 degrees
,'espectively. Figures 9c and 9d are for testconditions of M=0.22, Rc_5.6 million, and _---40 ° and
J_5°, respectively.
Space limitations inside the 1/16 t_ scale NTF model
due to the flow shield around the balance, and the
pressure instrumentation prohibited the use of an on
board accelerometer to measure model angle-of-attack. These angles were measured using an
arcsector mounted accelerometer package correctedfor sting bending using the balance loads and support
sting deflection sensitivities. These angle of attackmeasurements had an estimated accuracy of _0.1 °.Further information on the tunnel instrumentation,
data recording, and the data reduction algorithms is
provided in reference 6. The data herein were not
corrected for wall interference, support tare andinterference, and tunnel upflow.
Test Conditions
The NTF test program was designed to match Mach
number, chord Reynolds number, and angle of attack
with existing X-29A high alpha flight-test data. Thetest had a Mach number range from 0.22 to 0.25 atReynolds numbers based on mean aerodynamic
chord ranging from 0.7 to 6.8 million, and an angle-of-attack range from 28 to 68 degrees. A limited set
of data were acquired at 0.6 Mach number, Reynolds
numbers ranging from 1.6 to 8.3 million, and anangle of attack range from 28 to 42 degrees. The low
Reynolds number testing (Rc <_3.3 million) in air wasconducted at total pressures ranging fromapproximately 16 to 75 psia, and total temperaturesranging from 75 to 100°F. The high Reynolds
number testing (Rc > 5 million) in gaseous nitrogen
had total pressures ranging from approximately 30 to85 psia, with total temperatures ranging from -55 to-200°F. Overall the dynamic pressure ranged from
approximately 70 to 800 psf.
Results and Discussion
Tunnel to Flight Pressure Data ComparisonA comparison of the forebody pressure distributions
obtained from the NTF and flight is given in figures9 through 11. These data are plotted as the
coefficient of pressure (Cp) versus radial forebody
location (0) in degrees. Once again 0° represents thewindward side of the fuselage, 90 ° is the starboardside, and 180° is the leeward side. All the data in
3.0
2.5
2.0
1.5
1.0
-0.5
0.0
0.5
1.0 0
Flight M---0.250 Rc---6,83 x 10 e Alpha--30.1 de 9
,, NTF M=0.245 Rc--6.70 x 106 Alpha=30,0 deg
lm.
@=0"
x;I = 0,136
,,,,, ,, tit L_, _ J J ,I, t,, ,J J ,I, , : , t t, t I
90 180 270 360
o
9a) _ = 30 °
-3,0
-2.5
-2.0
-1.5
L"-l.0
-0.5
0,0
0.5
1.0
Right M=0.250 Rc--6.61x10 _ Alpha=34.9deg
-, NTF M--0.246 Rc=6,50x 106 Alpha=35.1 deg
_.(.... . ._ •
_, @=0"
0,136
C_ 'a.,
/ ',
I I I t I I i I [ I I I ' = ' ' = I J : _ + t : _ + I t , , t h J , , Ig0 180 270 360
o
9b) o¢,=35 o
Figure 9: NTF to Flight Forebody PressureDistributions for 30°< a < 45 ° at x/1=0.136
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American Institute of Aerona.ttics and Astronautics
-3.0
-2.5
-2.0
-1.5
_-1.0
-0.5
0.0
0.5
1.0(_
• Right M=0.220 Rc=5.94x106 AJpha=39.7deg
,. NTF M=0.214 Rc=5.84 x 106 Alpha--39.6 deg
1=).
90°: i -i270°
0=0"
x,'l=0.136
i i i t J _ _ ,i, _ : J _ t, , I, J k_l J i I I I J a LL_ J I I
90 180 270 360o
9c) e_ = 40 °
-3.0
-2.5
-2.0
-1.5
0.rO -1.0
-0.5
0.0
0.5
1.0
M=O.230 Rc=5.67x 10 e Alpha=49.7deg
r,/r=0.224 Rc=5.53x 10 s AJpha=,49.6deg
leO"
,oI 12,o
/
/
90 180 270 360o
10a) cc = 50 °
-3.0
-2.5
-2.0
-1.5
r,D_-1.0
-0.5
0.0
0.5
1.0
Flight M=0.220 Rc=5.55 x 106 Alpha=45.2 deg
NTF M=0.215 Rc--5.39x10 e Alpha=45.1deg
16o"
/
//
-._.,
90 180 270 360e
9d) cc ==45 °
Figure 9: Concluded
The primary suction peaks at 0=70 ° and 290 ° indicatethe local acceleration of the attached flow around this
highly curved region of the forebody surface. After
reaching the maximum forebody width, the flowbegins to decelerate as it approaches the leeward side
of the forebody. This deceleration continues until the
flow separates at 0=I 10 ° and 250 °. Finally, the
effects of the vortices due to the forebody are noted
by the secondary suction peaks at 0_ 150 ° and 210 °.This type of pressure distribution is typical of thatseen in previous forebody studies. 7m
An approximate 10 ° off set exists between the NTF
and the flight data on the starboard suction peak at
-3.0
-2.5
-2.0
-1.5
(D_-1.0
-0.5
0.0
0.5
1.0
+ FlightNTF
M--0.230 Rc=5.41 x 10 _ Alpha=54.7deg
M--0.226 Rc=5.31 x 10 * Alpha---.54.6 deg
/ 1Bo.
270*/r
0=O'
, _ _ , , , , , I , ..... XtI_ 0"136 t t _ I I90 180 270 360
10b) c_==55 °
Figure 10: NTF to Flight Forebody PressureDistributions 50 ° < ¢¢< 66 °
0"_70°. Since this off set remains fairly constant and
exists in all of the forebody pressure data examined,it is most likely attributed to a slight geometric
difference in the forebody cross-sectional geometrybetween the 1/16 _ scale NTF model and the X-29A
research airplane. As expected, all the pressuredistributions remain fairly symmetric in this alpha
range (30 ° < ct < 45°), and generally increase with
angle of attack. Overall there is a good correlationbetween the NTF and flight forebody pressuredistributions for angles of attack from 30 to 45degrees.
Figures 10a through 10d have test conditions of
M=0.23, Rc=5.4 million, and _z=50 °, 55 °, 59 °, and
66 °, respectively. There is still reasonable agreement
American Institute of Aeronautics and Astronautics
-3.0
-2.5
-2.0
-1.5
(D_-1.0
-0.5
0.0
0.5
1.0
• _qi_t_-, NTF
(: ',,
M---0.230 Rc--5.4.4x106 Alpha=Sg.0cleg
M=0.226 Rc--5.35x108 Alpha--59.1 deg
h¸
8=o o _;
9O 180 270 360
e
lOc) o_= 59 °
-3.0
-2.5
-2.0
-1.5
O_-1.0
-0.5
0,0
0.5
1.00
+ Flight
....... _ ........ NTF
/:
/
/
/
/
. , i t
M=0.220 Rc--5.18 x 10 6 Alpha--66.2 deg
M=0.215 Rc=5.10x 106 Alpha=66.2deg
_'i ..... ! -_270'
8 = O" :',,
x/I : 0,136
90 180 270 360
e
lOd) a = 66 °
Figure 10: Concluded
between the NTF data and the flight data at a-=50 °,but for angles of attack above 50 ° there is an
appreciable difference between the two pressure
distributions. At ct_55 ° a distinct asymmetrydevelops between the forebody vortices in the flight
data as indicated by the asymmetric secondary
suction peaks. The starboard vortex at 0_140 ° liftsaway from the surface while the port vortex at
0=210 ° shifts closer to the forebody causing a highersecondary suction peak under the vortex core. The
proximity of the port vortex to the forebody also
influences the primary suction peak at 0=290 °, andultimately results in a nose left yawing moment for
the research airplane. The pressure distribution for
the NTF model is more symmetric than the flight dataat this same angle of attack with only a slight nose
tight yawing moment indication. The secondary:uction peak under the port forebody vortex fox"the
NTF data is less pronounced here than it has been atlhe lower angles of attack.
•kt cz_-59" the asymmetry between the forebody,'ortices in the flight pressure distribution is more
i_ronounced, and again a pressure distributionassociated with the nose left tending yawing moment
:'.sobserved. The NTF data at ot_59" again is a more
:..ymmetric than that of the flight data with only a
:.;light tendency toward a nose fight yawing moment.
'rhe flight pressure distribution for ¢z=66 ° indicate a,:hange in asymmetry resulting in a nose right yawing
_noment for the research airplane, which is typical forvery sensitive high Reynolds number forebody apexflow fields. ''s'_ The NTF data at o_-_66° maintain
,:haracteristics similar to c_59 °, and unlike the flightJata did not experience a change in yawing momentdirection.
Overall these differences in the pressure distributions
:_etween the NTF and flight are most likely caused by
_he differences in both the boundary layer states, andthe geometric modeling of the forebody apex, noseboom, and nose strakes. The differences in the
boundary layers between the research airplane andthe 1/16 '_ scale NTF model may be attributed to
differences in the surface roughness between the twolest articles. The NTF model had a very smooth
surface finish (approximately 10 microinches), while
the research airplane had longitudinal gaps and stepsin the forebody due to instrumentation access panelsthat were located forward of the x/i=O. 136 pressure
row. Other external equipment on the research
airplane that could have affected the forebody flowespecially at the higher angles of attack include an
antenna, as well as the three angle of attack and oneangle of sideslip vanes mounted on the nose boom.None of these access panels or other equipment wasmodeled on the 1/16 t_ scale NTF test article. When it
is important to match high angle-of-attack flight
conditions for this type of forebody flow field, then itis necessary to consider even the smallest geometric
differences that may cause an asymmetry in the flow.
A source of error that may also contribute to the
discrepancies observed between the NTF and flightdata for c_ > 50 ° would be the wall interference
associated with using knuckle position #3 on the high
alpha sting. For this test, knuckle position #3 placedthe model in the closest proximity to the walls and
makes the pressure distributions more susceptible towall interference.
7
American Institute of Aeronautics and Astronautics
-3.0
-2.5
-2.0
-1.5
O%1.0
-0.5
0.0
0.5
1.0(_
-3.0
-2.5
Right M=0.600 Rc=8.03xl0 _ Alpt_-30.3deg
NTF M=0.598 Rc=7.g7x10 *¢ Alpha=30.2deg
f.----...,
\ i "_. //'
0=o °xll : 0.136
-2.0
-1.5
(._-1.0
-0.5
0.0
0.5
1.0
JJ=rklellJi]JJililllll_k:l_JJ11111190 190 270 360
e
1 la) a = 30 °
Right M=0.600 _m_ X 1_ _.0 _
....• Nil= M=0.598Rc=8.33x106AIl:m=34.9deg
150"
0=0".t/I : 0.136
_ _ t .... I ........ I, , , ..... I ........ I90 180 270 360
e
llb) (x ,,,,35 °
Figure 11:M--0.6 NTF to Flight ForebodyPressure Distributions 30 ° < a < 40 °
A limited set of higher Reynolds number data wereobtained at M_-_-0.6during the NTF test for
comparison with flight. Figures lla through liehave test conditions of M_-_O.6,Rc_8.2 million, and
ct=30 °, 35 °, 40 ° respectively. These pressuredistributions exhibit similar characteristics as seen in
the previous figures for flight Reynolds numbers of 5to 6 million at lower Mach numbers, however there is
a larger offset between the NTF and the flight data inthe vicinity of the forebody vortices. This offset
between the pressure distributions appears to remainfairly consistent over the limited angle of attackrange shown in figure 11, and would most likely beattributable to the differences in the state of the
boundary layers affecting separation locations on theleeward side of the forebody.
-3.0
-2.5
-2.0
-1.5
(.D_-1.0
-0.5
• - Right M=0.600 P,c=8.31x106 All:fna=_.2deg
,:. NTF M=0.598 P,c=8.26x10 s Al#'_,=,39.2deg
-3.0
-2.5
-2.0
-1,5
(0%1.0
-0.5
0.0
+ NIF M=0.216 Rc=0.74x106 ,.N#ka=45.2deg
NTF M=0.217 Rc=1.91x106 Alptk_45.3deg
,: NrF M=0.217 Rc=3.18x106 Alpha=,45.3deg
+ NTF M=0.215 F_=5.39x10 s _.4cleg
Kei ..... : i2_*
' _ Ill L _ J J ' '' LI' ] ' ' d k [ ]1] ....... I
1.0_ .... 90 180 270 360
o
12a) (x = 45 °
Figure 12: Reynolds Number Effects on ForehodyPressures (x = 45 ° and 66 o
Reynolds Number Effects on the Forebodv Flow
A unique advantage of testing in the NTF was theability to study the X-29 over a large range of
Reynolds numbers, Figure 12 shows forebodypressure data for the NTF model at chord Reynoldsnumbers ranging from 0.7 to 5.4 million. Figures12a and 12b have a test Mach number of
approximately 0.22 and angles of attack ofapproximately 45 ° and 66 ° respectively. All thesedata were obtained without fixing transition on the1/16 th scale NTF model. As shown in Figure 13 the
Reynolds numbers based on forebody diameter (RED)
in Lamont' s criteria range from a laminar boundarylayer state to a fully turbulent boundary layer state. _
American Institute of Aeronautics and Astronautics
AIAA-2003-0752
X-29 High Alpha Test in the National TransonicFacility (Invited)
Pamela J. Underwood, ViGYAN, Inc.
Hampton, VA 23666
Lewis R. Owens, Richard A. Wahls, and Susan Williams
NASA Langley Research Center
Hampton, VA 23681
41st AIAA Aerospace Sciences Meeting & Exhibit6- 9 January 2003
Reno, Nevada
For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics1801 Alexander Bell Drive, Suite 500, Reston, VA 20191-4344
AIAA-2003-0752
X-29 High Alpha Test in the National Transonic Facility
Pamela J. Underwood*
ViGYAN, Inc., Hampton, VA 23666
Lewis R. Owens _, Richard A. Wahls', Susan Williams _
NASA Langley Research Center, Hampton, VA 23681
Abstract Introduction
This paper describes the X-29A research program at
the National Transonic Facility. This wind tunnel
test leveraged the X-29A high alpha flight test
program by enabling ground-to-flight correlation
studies with an emphasis on Reynolds number
effects. The background and objectives of this test
program, as well as the comparison of high Reynolds
number wind tunnel data to X-29A flight test data are
presented. The effects of Reynolds number on the
forebody pressures at high angles of attack are also
presented. The purpose of this paper is to document
this test and serve as a reference for future ground-to-
flight correlation studies, and high angle-of-attack
investigations. Good ground-to-flight correlations
were observed for angles of attack up to 50 ° , and
Reynolds number effects were also observed.
Nomenclature
c mean aerodynamic chord, innormal force coefficientCI,I
Cr
Cv
DARPA
ESP
F
pressure coefficientside force coefficient
M
NTF National Transonic Facility
Pl
Rc
ReD
USAF
Defense Advanced Research
Projects Agency
electronically scanned pressure
degrees Fahrenheit
fuselage or body length, inMach number
total pressure, psi
Reynolds number based on mean
aerodynamic chord
Reynolds number based on
forebodv diameter
Wing reference area, in"
total temperature, °F
United States Air Force
axial distance from nose apex, in
angle of attack, deg
circumferential angle, deg
" Research Engineer, Member A1AA
i Aerospace Engineer. Flow Physics and Control Branch, SeniorMember AIAA
t Asst. Head. Configuration Aerodynamics Branch, AssociateFellow AIAA
Aerospace EngineerThis material is declared work of the U.S. Government and is not
subject to copyright protection in the United States.
A joint research program to investigate the high
angle-of-attack performance potential of the X-29A
forward swept wing fighter commenced in the
1980's. Primary partners in this joint program were
NASA, Grumman Aerospace, the United States Air
Force (USAF), and the Defense Advanced Research
Projects Agency (DARPA). During the high angle-
of-attack flight test, data were obtained to support
aerodynamics, flow visualization, control systems,
handling qualities, and maneuverability research. A
more extensive description of this flight program may
be found in reference 1. The National Transonic
Facility (NTF) X-29 High Alpha test was established
to augment this flight research program by enabling
ground-to-flight correlation studies with emphasis on
Reynolds number effects. Portions of the data from
this wind tunnel test have been previously published
in related articles. 2'3 The purpose of this paper is to
further document this test and serve as a reference for
future ground-to-flight correlation studies, and high
angle-of-attack investigations.
Aircraft Description
The X-29A is a single place research airplane with a
29.27 ° forward leading edge sweep wing, close
coupled variable incidence canards, and full span
dual hinged flaperons. 4 The Grumman Aerospace
designed X-29A is shown in figure 1. The forward
swept wing is an aeroelastically tailored composite
structure with a supercritical airfoil and a fixed
leading edge. A 75-inch long nose boom, and 24-
inch long nose strakes are positioned at the nose
apex. Side and bottom views of the nose apex are
shown in figure 2.
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Figure 1: Grumman Aerospace X-29A
Angle-of-allac_
Topview //_
_ , 180' ,
90o 270o :
e=o _ l_eO__ 10o" f: _,x/1:0.026 _ _ _ : !
9o.(.......i......./ ! .............,...............i2,0°, . ' _ _ :
o=o \ i / _! /
0:0 ox/I =0.201
X/I: 0.136 _,_
Figure 3: X-29A Research Aircraft ForebodyPressure Locations
The distribution of these static pressure rows is also
shown in figure 3, where 0 ° is the windward side of
the fuselage, 90 ° is the starboard side, and 180 ° is the
leeward side. In addition to these static pressure
orifices on the forebody, the research airplane was
also instrumented with three angle of attack vanes
and one angle of sideslip vane on the nose boom.
Facility Description
The NTF is a unique transonic wind tunnel designed
to conduct full-scale flight Reynolds number testing
through the use of high pressures and cryogenic
temperatures. This is a fan driven, closed circuit
wind tunnel with an 8.2-foot by 8.2-foot and 25-foot
long test section with a slotted ceiling and floor. A
planform view of the tunnel is provided in figure 4.
Figure 2: Side and Bottom View of Nose Apex
The nose boom tapered from a 0.88-inch diameter at
the tip to a 3.5-inch diameter at the nose apex, and
the nose strakes were 1.5 inches wide at the nose
apex and 2.5 inches wide at the downstream end.
As shown in figure 3 the research airplane was
instrumented with four circumferential rows of static
pressure orifices.
I_ 200 -- - _i
I[ Low-speed diffuser.. 19.7-dia fan ITurn3 : " _ ' I Turn2
1% ' _.?7---_--_r_
48.6 i ' 35.7 dta , 14.95:l_contractiono _
,0 ,aTurn 4 " L Screens _ High-speed diffuser
Cooling coil , 27-dia plenum 2.6 ° half-angle
± Wide-angle diffuser Slotted test section8.2 by 8.2
Figure 4: Planform View of the NTF
The tunnel operates using either dry air or gaseous
nitrogen as the test medium. During air operations
the tunnel pressure is used to control Reynolds
number, while in the cryogenic nitrogen mode the
tunnel temperature and pressure are used to control
Reynolds number. The NTF affords a test
2
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environment in which the Mach number and chord
Reynolds number are identical for the scale model inthe test section and the full-scale aircraft in flight.
The NTF is capable of an absolute pressure rangefrom 15 psia to 125 psia, a temperature range from
-320°F to 150°F, a Mach number range from 0.2 to
1.2, and a maximum Reynolds number of 146x106
per ft at Mach 1. Typical tests use a temperature
range from -250°F to ! 20°F. A more extensive
facility description can be found in reference 5.
NTF X-29 Test Program
The primary test objectives were to compare the NTF
high Reynolds number forebody pressure data to thedata obtained during the X-29A high alpha flight test,
and to assess the Reynolds number effects on theforebody flow at high angles of attack. The effect of
fixing transition on the forebody was also studied
during this test program.
1/16 th Scale X-29 Model
The NTF X-29 model is a 1/16 t_ scale representation
of the research airplane. All of the components of theX-29A research airplane were accurately scaled for
the NTF model except for the thickness of the nosestrakes. At 1/16 th scale the model nose strakes should
have been 0.0075 inches thick, but were actually 0.03inches thick due to NTF model strength
requirements. Pertinent model geometry is given in
figure 5.
The 1/16 thscale NTF model featured flow through
inlets positioned on either side of the forebody just
lc_rward of the canards that combined to form a single_;xhaust at the back of the model. A flow shield wasincluded to isolate the balance from the interior duct
flow in this model.
The contour tolerance of the wing, canard, and
_ertical tail was _+0.002 inches. The fuselage
torebody tolerance was __.0.004 inches, while thelemaining fuselage tolerance was approximately:,: 0.004 inches to _+0.006 inches. The model was
built of 18% nickel maraging steel (C type) with a
_,urface finish of approximately 10 microinchesRMS). The model was composed of separable
components to allow testing of multiple
_:onfigurations. The flaperons, aft body strakes,;-udder, and canards were all designed to be set atdiscrete angles. The 1/16 th scale NTF X-29 model is
_hown in figure 6 with all its control surface
uomponents. During this NTF test only the canard_mgle was varied. The model canard was designed toaccommodate five discrete angle positions (-20 °,
-25 ° , -30 ° , -35 °, -60°), and was set to match the flight
_est conditions as closely as possible.
Adjustable Control Surfaces
Flow Through _'///\
e = 5.41 in _/J
S -- 104.1 in2
Figure 5:1/16 _ Scale X-29 Model Geometry
20.4 in
Figure 6: NTF X-29 Model with Control Surfaces
A unique high alpha sting was used with the X-29model. This sting was designed to accommodate an
angle of attack range from -7 ° to 74 ° using threeprimary knuckle positions. A sketch of this high
alpha sting showing the three possible knucklepositions is given in Figure 7. As seen in this sketch,only the second knuckle position keeps the model
center of rotation aligned with the tunnel centerline,
and positions #1 and #3 would present the modelpositions most susceptible for wall interference.Figure 8 shows the 1/16 th scale X-29 model mounted
on the high alpha sting in the NTF test section.
American Institute of Aeronautics and Astronautics
_._, Position #3
_ ..... !r 1_
i Position______/#2/ -_F - f "__..
Position #1
Figure 7: General Arrangement of NTF High
Alpha Sting Depicting Three Knuckle Positions
(Not to Scale)
Figure 8:X-29 Model Mounted in NTF Test Section
(Knuckle Position #3)
Instrumentation and Measurement Corrections
The NTF model was instrumented with one
circumferential row of static pressure orifices on the
forebody at station x/l=0.136. This forebody location
was chosen to correspond with one of the X-29A
research airplane static pressure row locations (see
figure 3). The circumferential distribution of the
model static pressure orifices was also positioned to
match the research airplane as closely as possible.
Table 1 gives the circumferential pressure orifice
locations for both the X-29 research airplane and the
X-29 NTF model at x/l=0.136. Again, 0 ° is the
windward side of the fuselage, 90 ° is the starboardside, and 180 ° is the leeward side.
electronically scanned pressure (ESP) module with a
full-scale pressure range of_+ 15 psid. The quoted
(worst case) accuracy of the ESP module was
approximately 0.20 percent of full scale or _+0.030
psi (Cp variation _-0.06 at the lowest dynamic
pressure condition). For reference, the X-29A high
alpha flight test forebody pressure data were obtained
using _+ 1.5 psi differential pressure transducers with
an estimated accuracy of approximately _+0.007 psi. 4
The model aerodynamic force and moment data were
obtained with a six component unheated strain gage
balance. The balance maximum load capacity and
quoted accuracies are given in Table 2. The axial
force and moment data acquired were inconsistent,
and deemed corrupt. These data show signs of
interference on the balance most likely due to the
tightly packed instrumentation within the model.
Normal and side force data were less sensitive to this
adverse effect, and are presented herein. The main
objective was to compare forebody pressures with
available flight data.
Table 1: Circumferential Static Pressure Locations,
x/!=0.136
NTF Flight
0 (deg) 0 (de_:)24.2 n/a
47.6 n/a
59.6 60.0
n/a 66.0
72.0 72.0
77.9 78.0
83.9 n/a
89.8 90.0
95.0 95.0
99.9 100.0
103.7 105.0
107.7 108.0
114.1 111.1
119.9 120.0
125.9 126.0
n/a 129.0
131.9 132.0
rda 135.0
138.0 138.0
n/a 141.0
143.9 144.0
n/a 147.0
150.0 150,0
n/a 153.0
155.9 156.0
162.1 165.0
167.9 168.0
n/a 171.0
174.0 174.0
180.0 n/a
The NTF model forebody pressure data were
obtained through the use of one intemal 48 port
4
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NTF Flight
0 (deg) 0 (de_)185.9 186.0
n/a 189.0
191.9 192.0
197.7 195.0
203.6 204.0
n/a 207.0
209.8 210.0
rda 213.0
215.5 216.0
n/a 219.0
221.4 222.0
n/a 225.0
227.3 228.0
n/a 231.0
234.7 234.0
n/a 237.0
240.5 240.0
246.2 249.0
252.3 252.0
n/a 255.0
260.3 260.0
265.2 265.0
270.0 270.0
276.0 276.0
288.0 288.0
n/a 294.5
299.5 300.0
336. t n/a
360.0 n/a
Table 2: NTF Balance Load Capacity
Component Maximum Full-ScaleLoad Accuracy
Normal Force _ 2500 lb 0.10 %Axial Force _ 350 lb 0.26 %
_ 5000 in-lb 0. I 1%_+2500 in-lb 0.40 %
Pitching MomentRolling MomentYawing Moment
Side Force_+4000 in-lb 0.18 %_ 1000 lb 0.35 %
these figures were obtained without fixing transitionon the 1/16 th scale NTF model. Figures 9a and 9b
are for test conditions of 0.25 Mach number (M),
Reynolds number based on mean aerodynamic chord(Rc) of approximately 6.6 million, and angles of
attack (c_) of approximately 30 and 35 degreesrespectively. Figures 9c and 9d are for test
conditions of M_0.22, Rc_5.6 million, and _=40" and45 °, respectively.
Space limitations inside the 1/16 thscale NTF model
due to the flow shield around the balance, and the .3o
pressure instrumentation prohibited the use of an on
board accelerometer to measure model angle-of- .2.5attack. These angles were measured using an
arcsector mounted accelerometer package corrected .2.0
for sting bending using the balance loads and supportsting deflection sensitivities. These angle of attack .1.5measurements had an estimated accuracy of _+0.1°.Further information on the tunnel instrumentation, c3-1.0
data recording, and the data reduction algorithms is-0.5
provided in reference 6. The data herein were not
corrected for wall interference, support tare andinterference, and tunnel upflow. 00
0.5
1.0
Test Conditions
The NTF test program was designed to match Machnumber, chord Reynolds number, and angle of attack
with existing X-29A high alpha flight-test data. Thetest had a Mach number range from 0.22 to 0.25 at
Reynolds numbers based on mean aerodynamic
chord ranging from 0.7 to 6.8 million, and an angle-of-attack range from 28 to 68 degrees. A limited set 3.0
of data were acquired at 0.6 Mach number, Reynolds 2.5numbers ranging from 1.6 to 8.3 million, and an
angle of attack range from 28 to 42 degrees. The low 20Reynolds number testing (Rc < 3.3 million) in air wasconducted at total pressures ranging from 15
approximately 16 to 75 psia, and total temperaturesranging from 75 to 100°F. The high Reynolds G lo
number testing (Rc _>5 million) in gaseous nitrogenhad total pressures ranging from approximately 30 to o.585 psia, with total temperatures ranging from -55 to
-200°F. Overall the dynamic pressure ranged from 0.0approximately 70 to 800 psf.
3.5
Results and Discussion
• Flight M--0.250 Rc--6.83 x 10 _ Alpha--30.1 deg
. NTF M---0.245 Rc---6.70xl06 AJpha--<30.0deg
1.o-
!_ ............... i27o°
\ ,', j
e:0"
xTI : 0,136
, _ J J , i i . I i h I I i I I h I , i i J i , i i I , , , , , , , I
90 180 270 360e
--_l_ RightNTF
9a) _= 30 °
1.0
M=0.250 Rc=6.61 xl06 AJpha=34.gdegM=0.246 Rc=6.50x 10_ Alpha--35.1 deg
: t
.,( ',
/ "
/
Tunnel to Flight Pressure Data ComparisonA comparison of the forebody pressure distributions
obtained from the NTF and flight is given in figures9 through 11. These data are plotted as the
coefficient of pressure (Cp) versus radial forebody
location (0) in degrees. Once again 0° represents thewindward side of the fuselage, 90 ° is the starboardside, and 180 ° is the leeward side. All the data in
5
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, ,,,,,, _ I, _ J i _ L i i I t , J J _ i I liill I I I ill90 180 270 360
o
9b) ct = 35 °
Figure 9: NTF to Flight Forebody PressureDistributions for 30°< ct < 45 ° at x/!=0.136
-3.0
-2.5
-2.0
-1.5
o.
(..) -1 .(3
-0.5
0.0
0.5
1.0 0
t
?:"
I:
Right M=0.220 Rc=5.94x10 _ Alpha=39.7deg
NTF M=02.14 Rc=5.84x10 _ Alpha=39.6deg
18o"
\,, j
_I =0.136
',%--f
J J J I k J d I _ d J I I J i _ t ( J i i L _ J I , I L , ] _ L L i J I
90 180 270 360
e
9c) a " 40 °
-3.0
-2.5
-2.0
-1.5
f._%1.0
-0.5
0.0
0.5
1.0,
Right M=0.230 Rc=5.67x106 Alpha=49.7deg
_., NIT M=0.224 Rc=5.53 x 106 Alpha=49.6 deg
,0S 1,,o
t _t=0 °j x/t= 0.136
/
........ I ........ I ....... _lhl _ _ r i i _190 180 270 360
o
10a) (_ = 50 °
-3.0
-2.5
-2,0
-1.5
o.-1.0'
-0.5
0.0
0.5
1.0
Flight M=0.220 Rc=5.55x10 e Alpha=45.2deg
•"_ -- NTF M=0.215 Rc--5.39x 106 Alpha--45.1 deg
Figure 9: Concluded
The primary suction peaks at 0=70 ° and 290 ° indicatethe local acceleration of the attached flow around this
highly curved region of the forebody surface. Afterreaching the maximum forebody width, the flow
begins to decelerate as it approaches the leeward sideof the forebody. This deceleration continues until the
flow separates at 0=110 ° and 250 °. Finally, the
effects of the vortices due to the forebody are noted
by the secondary suction peaks at 0=_150 ° and 210 °.
This type of pressure distribution is typical of that7_0
seen in previous forebody studies. "
An approximate 10° off set exists between the NTFand the flight data on the starboard suction peak at
6
-3.0
-2.5
-2.0
-1.5
¢.._%1.0
-0.5
0.0
0.5
1.0
Flight M=0.230 Rc=5,41 x 106 Alpha=54.7de(j
NTF M=0.226 Rc=5.31 x 106 Alpha=54.6 deg
/ :
0=0"
........ j ...... ,_t--0.?_ ....... _ ........90 180 270 360
o
lOb) a ,_ 55 °
Figure 10: NTF to Flight Forebody PressureDistributions 50 ° < (x < 66 °
0=70 °. Since this off set remains fairly constant andexists in all of the forebody pressure data examined,
it is most likely attributed to a slight geometricdifference in the forebody cross-sectional geometrybetween the 1/16 th scale NTF model and the X-29A
research airplane. As expected, all the pressure
distributions remain fairly symmetric in this alpha
range (30 ° < a < 45°), and generally increase with
angle of attack. Overall there is a good correlationbetween the NTF and flight forebody pressuredistributions for angles of attack from 30 to 45
degrees.
Figures 10a through 10d have test conditions of
M-_-O.23, Rc_5.4 million, and c_=50°, 55°, 59 °, and
66°, respectively. There is still reasonable agreement
American Institute of Aeronautics and Astronautics
-3.0
-2.5
-2,0
-1,5
r,_ - 1.0
-0,5
0.0
0.5
1.0 o
• Right M=0.230 Rc=5.44X10 _ Alpha=59.0deg
(, NTF M--0.226 Rc=5.35 x 106 Alpha=59.1 deg
10c) (_ = 59 °
+ Flight
,!:..... N"rF
90
M=0.220 Rc--5.18 x 106 Alpha=.66.2 deg
M=0.215 Rc--5,10x 106 Alpha--66.2deg
x,'l = 0.136
o
lOd) _ = 66 °
270
Figure 10: Concluded
between the NTF data and the flight data at _=50 °
but for angles of attack above 50° there is an
appreciable difference between the two pressure
distributions. At (x=55 ° a distinct asymmetrydevelops between the forebody vortices in the flight
data as indicated by the asymmetric secondary
suction peaks. The starboard vortex at 0"_140 ° liftsaway from the surface while the port vortex at
0_210 ° shifts closer to the forebody causing a higher
secondary suction peak under the vortex core. Theproximity of the port vortex to the forebody also
influences the primary suction peak at 0=290 °, andultimately results in a nose left yawing moment for
the research airplane. The pressure distribution forthe NTF model is more symmetric than the flight data
at this same angle of attack with only a slight nose
36O
tight yawing moment indication. The secondary
suction peak under the port forebody vortex for the
NTF data is less pronounced here than it has been atlae lower angles of attack.
At ff_-59 ° the asymmetry between the forebody
,ortices in the flight pressure distribution is morepronounced, and again a pressure distribution
associated with the nose left tending yawing moment
is observed. The NTF data at c_59" again is a more
_ymmetric than that of the flight data with only a,qight tendency toward a nose right yawing moment.
"['he flight pressure distribution for ot=66 ° indicate a
change in asymmetry resulting in a nose right yawing
moment for the research airplane, which is typical forvery sensitive high Reynolds number forebody apexflow fields, z'8`_ The NTF data at ct=66 ° maintain
characteristics similar to e¢=59 °, and unlike the flight
data did not experience a change in yawing momentdirection.
()verall these differences in the pressure distributions
between the NTF and flight are most likely caused bythe differences in both the boundary layer states, and
the geometric modeling of the forebody apex, noseboom, and nose strakes. The differences in the
boundary layers between the research airplane andthe 1/16 th scale NTF model may be attributed to
differences in the surface roughness between the twotest articles. The NTF model had a very smooth
surface finish (approximately 10 microinches), whilethe research airplane had longitudinal gaps and steps
in the forebody due to instrumentation access panelsthat were located forward of the x/l=O. 136 pressure
row. Other external equipment on the research
airplane that could have affected the forebody flowespecially at the higher angles of attack include anantenna, as well as the three angle of attack and one
angle of sideslip vanes mounted on the nose boom.None of these access panels or other equipment wasmodeled on the 1/16 _ scale NTF test article. When it
is important to match high angle-of-attack flight
conditions for this type of forebody flow field, then itis necessary to consider even the smallest geometric
differences that may cause an asymmetry in the flow.
A source of error that may also contribute to thediscrepancies observed between the NTF and flightdata for cx > 50 ° would be the wall interference
associated with using knuckle position #3 on the high
alpha sting. For this test, knuckle position #3 placedthe model in the closest proximity to the walls andmakes the pressure distributions more susceptible towall interference.
7
American Institute of Aeronautics and Astronautics
.3.0 -
-2.5
-2.0
-1.5
(_%1.0
-0.5
O0
0.5
1,0(3
-3+0
-2.5
-2.0
-1.5
o.L_ -1.0
-0.5
0.0
0+5
1.0_
Right M=0.600 Rc=8.03xl06 Alpha,=30.3dog
NTF M=0.598 Rc=7.97x10 + Alpha=30.2deg
180'
\ :
tt=0'
xJl = 0.136
illlllllllllllJRLIIiililll|l+;l_lt,I
1_ 2_ 360e
1 la) a = 30 °
+ Right Ivl=0.600 Ft_=8.42x10 + ,411:tmm35.0deg
• , NTF M=0.596 Rc=8.33x106 Alpt',a=34.9deg
lao+
" + i.... '
x,l =0.136
, J , i , < i | i , J i ,L= t | t , I li il i I J i ,h h J t J I
90 180 270 360e
IIb) ¢x= 35 °
Figure 11:M--0.6 NTF to Flight ForebodyPressure Distributions 30 ° < ¢x< 40 °
A limited set of higher Reynolds number data were
obtained at M-_.0.6 during the NTF test for
comparison with flight. Figures lla through llchave test conditions of M-_0.6, Rc_8.2 million, and
a_30 °. 35 °, 40 ° respectively. These pressuredistributions exhibit similar characteristics as seen in
the previous figures for flight Reynolds numbers of 5to 6 million at lower Mach numbers, however there is
a larger offset between the NTF and the flight data inthe vicinity of the forebody vortices. This offset
between the pressure distributions appears to remain
fairly consistent over the limited angle of attackrange shown in figure 11, and would most likely beattributable to the differences in the state of the
boundary layers affecting separation locations on the
leeward side of the forebody.
-3.0
-2.5
-2.0
-1.5
(.)_-1.0
-0.5
0+0
0.5
1.0
-3.0
-2.5
-2.0
-1.5
C._- 1.0
• Rigid M=O.600 Rc=8.31x10 + _.2deg
.... NFF M=0.596 Rc=8.26x10 + /Wpha=39.2deg
180"
lo*; .............. ;270
tt=O*
x,l=0.136
+'+,,,iliili,_J,l=_ .... il+_++++++l
1_ 27O 360
e
llc) a =40 °
----.1-- NTF M=O.216 Rc-..-0.74x10 + ,41I:tta=45.2deg
--_'-.+--- NTF 171=0.217 Rc=1.91x10 + Alpha:=45.3deg
,:+_ NFF IV1=0217 Iqc-=3.18x106 Alptmr::45.3deg
NTI= M=0.215P,:5.39x10"._plm=45.4deglil0 •
90"; ...... :270
.-0.5
0.0
0.5
1,0dl t I .... IIIliiiiiili i,,,,_|, ,illili I90 180 270 360
0
12a) or.= 45 °
Figure 12: Reynolds Number Effects on ForebodyPressures ix _ 45 ° and 66 o
Reynolds Number Effects on the Forebody FlowA unique advantage of testing in the NTF was the
ability to study the X-29 over a large range of
Reynolds numbers. Figure 12 shows forebodypressure data for the NTF model at chord Reynoldsnumbers ranging from 0.7 to 5.4 million. Figures12a and 12b have a test Mach number of
approximately 0.22 and angles of attack of
approximately 45 ° and 66 ° respectively. All thesedata were obtained without fixing transition on the1/16 thscale NTF model. As shown in Figure 13 the
Reynolds numbers based on forebody diameter (RED)
in Lamont's criteria range from a laminar boundary
layer state to a fully turbulent boundary layer state. 7
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• NTF IVL-0.217 P,c=0.68x106 Alpha=66.0deg
-3.0 ,; NTF M=0.217 Rc=l.90xl0 _ Alpha=65.9deg
,,, NTF M=0.218 Rc--3.18x10 _ Alpha=65.9deg
_+ NTF M=0.217 Rc--5.10xl0 r" Alpha=66.0deg-2.5 _ 1_o
t '_ _....... ! i _ \-1.5 C,< _-__-J" ., .
d-,o i;
,,' ,_
0.0
(1.5
10 I i I i i I i | I I I I I i I I ] I i , ..... | , i d L L . -_' 90 180 270 360
o
12b) a = 66 °
Figure 12: Concluded
g0¢-
¢
5 J
4 FT FT FTFTFTFTFTFTFTFTFT/'r T.j,
3 FT FE_ FTFTFTFTT"rlL_FT?FT?T TT(I_ULLY TURBULENT} •
2 FT F_T .FTFTFTFTFT/T_TT T T
FT _ FTFTFTTT TiT T T T
1 / ( TRAIII$1TIONAL I
FT I_T FTT T T T T_T T T T
_T T T.5 |
T _ T T T T T TIT T T T
.2 _L L L L L L L L k
( L_INAR )• & _ L I ,I J • • _-
0 10 20 30 40 50 60 7C 80 90
ANGLEOFATTACK,a. de_
Figure 13: Lamont's Classification of ThreeMainFlow Regimes 7
For reference, the ratio of the NTF model forebody
diameter (x/l=0.136) to the mean aerodynamic chordis approximately 2/5 or 0.4, and the approximate
range of ReD covered in this test is highlighted infigure 13.
c,ccurs at 0=90 ° and 270 '_. The higher Reynolds
number data in this figure exhibit a transitionalboundary layer characterized by the presence of a
separation bubble at 0-_ 100 ° and 260 °.
A distinct difference in the forebody flow is noted
when comparing the lowest Reynolds number
pressure distributions (Re=0.7 million or RED=0.28
million) for ct,_45 ° and 66 °. At et-_45° the forebodyflow exhibits more of a transitional boundary layerc:haracter while the data at cx=66 ° indicate a more
laminar boundary layer state. This demonstrates thelower critical Reynolds number boundary (betweenthe L and T flow regimes) variation with angle ofattack shown in figure 13. 7J2
These Reynolds number effects can also be detected
in the normal (CN) and side (Cv) force data. Figures14 and 15 show the CN and Cv data for the same test
conditions as figures 12a and 12b, respectively. The
moderate angle-of-attack data shown in figure 14does not indicate a significant Reynolds number
_;ffect on CN or C¥. All the pressure data for cz_-45"
exhibit a transitional to fully turbulent boundary layer
state, and as expected the variations in CN and Cvwith Pc- are minimal. 8J: The higher angle of attack
data shown in figure 15 reveal more variation CN and
Cv for the Rc=0.7 million condition. This higher
force data was expected since the pressure
distribution for o_66 ° had a more laminar boundarylayer characteristic._°
L00
As expected, the secondary suction peaks at 0= 140 °and 210 ° due to the forebody vortices are most
prevalent at the lowest and the highest Reynoldsnumbers. There are only small differences between
the higher Reynolds number data (Rc>1.9 million) infigure 12b. The lowest Reynolds number data(Rc=0.7 million) was fundamentally different at
c_66 °. The pressure distribution for this Reynoldsnumber resembles more of a laminar flow field since
there is little pressure recovery before separation
2.70
2.40
z
2.10
180
+ NTF M=0.217 Rc=0.67x10 e
C- NTF M=0,217 Rc=1.91 x 10s...._s-.... NTF M=0,216 Rc=3.18 x 10e
+ NTF M=0.214 Rc=5.35 x 10_
'.5o,2.... A .... 2,, ,'5.... 2o.... 2,.... ,'8c(
14a) Normal Force (CN) Variation
Figure 14: Reynolds Number Effects on CN and
Cy for _ = 45 °
9
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0.20
0.10
(j_ 0.00
-0.10
-_ NTF M=0.217 Rc=0.67x 104
:7 NTF M=0.217 Rc=1.91 x 106
,: NTF M--0.216 Rc=3.18x106
_ NTF M=0,214 Rc=5.35x10 _
0.20 -
0.10
C_ 0.00
-0.1_
-0.204'' _'lJ'JIl_l Lllll I hitlill2 43 44 45 46 47 .... 481 -0.2(
O_
14b) Side Force (Cy) Variation
t NTF M---0.217 Rc---0.68 x 106
-, NTF M=0.217 Rc=1.90 x 106
...., NTF M=0.217 Rc--3.18x 106
_ NTF M--'0.215 Rc=5.11 x 106
JiJiIttl,I .... I,,,biLkl,I .... I .... I .... I12 63 64 65 66 67 68 69 70
15b) Side Force (Cy) Variation
3.00,
2.70
2.40
d
2.10
1.80
1.5oB'2
Figure 14: Concluded
_-- NTF
.... _:_ NTF.... NTF
NTF
M=0.217 Rc=0.68 x 10 _
M=0.217 Rc=1.90 x 10 _M=0.217 Rc=3.18x 10 eM=0.215 Rc=5.11 x 106
63 64 65 66 67 68 69 70_E
15at Normal Force (CN) Variation
Figure 15: Reynolds Number Effects on CN and
Cy for a = 66 °
-3.0
-2.5
-2.0
-1.5
(_-1.0
-0.5
NrF k-t=0217 Rc=0.68x10 e Npha=45.4deg C_,-:-:-_#1
,; NTF M=0.217 Rc=0.68x10' /_pha=45.3deg Grit#'2
, NTF M=0.217 FIc=0.68x10 B AIioha=45.3deg NoGrit
NTF M=0.215 Rc=5.39x10 ° Alpha=45.1deg N_G_t
mo-
0.5
1.0 .... 90 180 270 360
e
Figure 16: Effect of Transition on ForebodyPressures at ¢x - 45°
Effect of Transition on the ForebodyThe effect of two different fixed transition patternson the NTF model forebody was also studied during
this test program. The purpose of this study was todetermine if better tunnel to flight correlation could
be attained through the use of fixed transition. Twingrit strips were applied starting at the end of the nose
strakes and extending approximately 7 inches tordl_-_.0.23. Transition pattern #1 was a band of#80
carborundum grit that had a constant width ofapproximately 0.25 inches. Transition pattern #2 was
also a band of #80 carborundum grit that varied inwidth from 0.25 inches wide at the nose strake to
approximately 1.0 inch wide at x/1=0.23. Figure 16shows forebody pressure distributions for both
transition patterns for M_0.22, Rc_.7 million, and
10
American Institute of Aeronautics and Astronautics
_-45 '_. For reference, transition free data at both the
low and high Reynolds number conditions is also
included in this figure. Note that at this angle ofattack the low and high Reynolds number free
transition data match reasonably well without anyforced transition. The data obtained from transition
pattern #1 resemble a fully turbulent pressuredistribution. The vortices due to the forebody are
more prominent for transition pattern #2. Both
transition patterns are reasonably symmetricanalogous to the NTF flight Reynolds number data at
this angle of attack. However, transition pattern #1appears to more closely simulate the high Reynoldsnumber condition.
-3.0
-2.5
M=0,217 Rc=0.68x106 Alpha--66.1 deg Grit#1M=0,218 Rc=O68 x 10_ Alpha=66.0 deg Grit #2
M---0.217 Rc=-068 x 106 Alpha----66.0dsg No GdtM--0217 Rc=5.10 x 106 Alpha---66.0 deg No Grit
is0o
t t90"I " ;270
j.... JJ
,_=0•x_l=0.136
1.00
Figure 17:
gO 180 270o
Effect of Transition on ForebodyPressures at cz = 66 °
360
Figure 17 shows forebody pressure distributions for
both transition patterns for M_0.22, Re=0.7 million,
and ct_-66 °. Again, transition pattern #1 resembles afully turbulent pressure distribution, and the effect of
the forebody vortices is most prominent for transitionpattern #2. Both transition patterns eliminate the
laminar flow field observed in the low Reynoldsnumber transition free pressure distribution at this
angle-of-attack. The high suction peaks at 0=70 ° and
290 ° for the NTF flight Reynolds number data are notmatched by either of the fixed transition patterns,
although slight asymmetries are observed in bothtransition pattern pressure distributions analogous to
the flight Reynolds number data. Fixed transitiondata were only obtained at the lowest Reynolds
number condition, Rc_.7 million, during this test
program.
Since the time of this test additional research has
been performed providing additional insight into
gritting strategies for high angle-of-attackinvestigations. _3Twin grit strips on the model
forebody are still prefen'ed, but the width of thesestrips is now recommended to be approximately O. 13inches. A constant width grit pattern is
recommended. It would be interesting to test thisnew transition pattern on the 1/16 _ scale X-29 model
at flight Reynolds number, Rc=5 million, for anglesof attack greater than 50 ° to see if an asymmetric
forebody flow field develops similar to thoseobserved in the X-29A flight data. It would also be
interesting to test this new transition pattern at all test
Reynolds numbers, not just flight, to determine if thenew pattern actually makes the low Reynolds
numbers better resemble the flight pressuredistributions.
Conclusion
Results from the NTF X-29 High Alpha test have
been presented. The NTF high Reynolds number
forebody pressure data and the X-29A flight test data
showed good correlation up to ct_50 °. For angles ofattack above 50 °, the flight pressure distributions
become asymmetric and do not correlate as well withthe high Reynolds number NTF data. The
differences in the pressure distributions wereattributed to a difference in the boundary layer statesbetween the NTF model and the X-29A research
airplane. The difference in the boundary layer states
is most likely caused by a difference in the surfaceroughness between the two test articles, and the
external equipment on the X-29A research airplaneforebody and nose boom that was not modeled on the1/16 th scale NTF model. The wall interference
associated with using knuckle position #3 on the NTF
X-29 high alpha sting may also contribute to thediscrepancies between the tunnel to flight pressuredistributions for angles of attack above 50°. The
Reynolds number effects on the NTF model forebodypressures for moderate and high angles of attack were
also presented. The lowest Reynolds number data
(Rc=0.7 million) at _=66 ° showed a laminar flowfield which was substantially different from the
higher Reynolds number (Rc > 1.9 million) pressuredistributions that exhibited more of a transitional
boundary layer characteristic. Fixing transition onthe NTF model forebody for the lowest test Reynolds
number condition improved the correlation to thehigher NTF Reynolds number data, but still showed
some fundamental differences with flight Reynoldsnumber pressure distribution.
11American Institute of Aeronautics and Astronautics
1)
2)
3)
4)
5)
References
Moore, M., Frei, D., X-29 Forward SweptWing Aerodynamic Overview, AIAA-83-1834, July 1983.
Fisher, D., Cobleigh, B., Banks, D., Hall, R.,
and Wahls, R., Reynolds Number Effects atHigh Angles of Attack, NASA TP-1998-206553, June1998.
Luckring, J.M., An Overview of NationalTransonic Facility Investigations for High
Pe_orrnance Militat T Aerodynamics,A1AA-2001-0906, January 2001.
Fisher, D., Richwine, D., Landers, S.,
Correlation of Forebody Pressures andAircraft Yawing Moments on the X-29A
Aircraft at High Angles of Attack, NASATM-4417, November 1992.
Fuller, D., Guide for Users of the National
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Moskovitz, C., Hall, R., DeJarnette, F.,
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85-1799, August 1985.
Hall, R., Erickson, G., Fox, C., Banks, D.,
Fisher, D., Evaluation of Gritth_g Strategiesfor High Angle of Attack Using Wind Tunnel
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Foster, J., and Adcock, J., User's Guide for
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Lamont, P.J., Pressures Around an lnclined
Ogive Cylinder with Laminar, Transitional,and Turbulent Separation, AIAA-80-1556R,March 1982.
Polhamus, E., A Review of Some Reynolds
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Roos, F.W., and Kegelman, J.T.,
Aerodynamic Characteristics of ThreeGeneric Forebodies at High Angles of
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10) Owens, L., Hemsch, M., Popernack, T.,
Reynolds Number Effects on AdvancedSlender Forebodies for Angles of Attack up
to 27 ° at Mach 0.2, NASA TP-3493, August1994.
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American Institute of Aeronautics and Astronautics
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