x-29 high alpha test in the national transonic facility ...p! rc red usaf total pressure, psi...

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.

1

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

2


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.

3


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

4


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

5

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


-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


-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. _


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.

1


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


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.


_._, 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


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


, ,,,,,, _ 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


-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


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


.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


• 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 °


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


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


_-- 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


_-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

Transonic Facility, NASA TM-83124, 1981.

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13)

Moskovitz, C., Hall, R., DeJarnette, F.,

Combined Effects of Nose Blunmess andSurface Perturbations on Asymmetric FlowPast Slender Bodies, Journal of Aircraft,Vol. 27, Number 10, October 1990.

Lamont, P.J., The Effect of ReynoldsNumber on Normal and Side Forces on

Ogive Cylinders at High Incidence, AIAA-

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

and Flight Test Data for the F/A-18, NASATP-1998-207670, May 1998.

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Foster, J., and Adcock, J., User's Guide for

the National Transonic Facility ResearchData System, NASA TM-110242, April1996.

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

Number Effects Related to Bodies at HighAngles of Attack, NASA CR-3809, August1984.

Roos, F.W., and Kegelman, J.T.,

Aerodynamic Characteristics of ThreeGeneric Forebodies at High Angles of

Attack, AIAA-91-0275, January 1991.

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.

12


x-29 high alpha test in the national transonic facility ...p! rc red usaf total pressure, psi...

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