Investigation of Reynolds Number Effects on a Generic Fighter Configuration in the National Transonic Facility

22 Pages·2002·0.93 MB·English

by  Tomek, W. G.

S_ -if-! C."._- JJ-J
AIAA-2002-0418
Investigation of Reynolds Number Effects on a
Generic Fighter Configuration in the
National Transonic Facility (Invited)
W. G. Tomek, R. M. Hall, R. A. Wahls, J. M. Luckring, and
L. R. Owens
NASA Langley Research Center
Hampton, Virginia
40th AIAA Aerospace Sciences Meeting & Exhibit
14-17 January 2002
Reno, Nevada
For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics
1801 Alexander Bell Drive, Suite 500, Reston, VA 20191-4344

AIAA-2002-0418
INVESTIGATION OF REYNOLDS NUMBER EFFECTS ON A
GENERIC FIGHTER CONFIGURATION
IN THE NATIONAL TRANSONIC FACILITY
W. G. Tomek 1, R. M. Hall 2, R. A. Wahls _, J. M. Luckring 4,and L. R. Owens s
Aerodynamics, Aerothermodynamics, and Acoustics Competency
NASA Langley Research Center
Hampton, Virginia
ABSTRACT INTRODUCTION
A wind tunnel test of a generic fighter Simulation of flight at full-scale Reynolds
configuration was tested in the National numbers is now available through the use of
Transonic Facility through a cooperative cryogenic wind tunnels, such as the National
agreement between NASA Langley Research Transonic Facility (NTF) at the NASA Langley
Center and McDonnell Douglas. The primary Research Center. Some of the initial models
purpose of the test was to assess Reynolds tested in this facility were those built to study the
number scale effects on a thin-wing, fighter-type Reynolds number effects on transport aircraft.
configuration up to full-scale flight conditions These models were composed of U.S. aircraft
(that is, Reynolds numbers of the order of 60 industry designed and fabricated wings mounted
million). The test included longitudinal and to a NASA supplied generic fuselage, denoted
lateral/directional studies at subsonic and as Pathfinder I (refs. 1 and 2). A similar need
transonic conditions across a range of Reynolds was identified for studying fighter aircraft
numbers from that available in conventional concepts using a generic, area-ruled fuselage,
wind tunnels to flight conditions. which could accommodate a variety of wing
Results are presented for three Mach planforms. NASA has denoted the fuselage for
numbers (0.6, 0.8, and 0.9) and three the configuration as Pathfinder I1. Further
configurations: 1) Fuselage / Wing, 2) Fuselage / general discussion of the Pathfinder II family of
Wing / Centerline Vertical Tail / Horizontal Tail, models can be found in references 1- 3.
and 3) Fuselage / Wing / Trailing-Edge The objectives of this wind tunnel
Extension / Twin Vertical Tails. Reynolds investigation, utilizing the Pathfinder II fuselage
number effects on the longitudinal aerodynamic with a McDonnell Douglas defined thin, fighter-
characteristics are presented herein. type wing, were to study the effects of Reynolds
number on a fighter-type configuration through
1Aerospace Engineer, Research Facilities Branch, Member,
model component build-up and with stability and
AIAA
control device deployment (ref. 4). The test was
2Aerospace Engineer, Configuration Aerodynamics Branch, conducted to provide a database of Reynolds
AssociateFellow,AIAA number effects, up to full scale, which could be
used to determine Reynolds number correlation
3Assistant Branch Head, Configuration Aerodynamic Branch,
trends, provide data for assessment of
AssociateFellow,AIAA
computational fluid dynamics (CFD) methods
4Aerospace Engineer, Configuration Aerodynamics Branch, including turbulence modeling, and validate
AssociateFellow,AIAA design and analysis methods.
This paper presents results from a
SAerospace Engineer, Flow Physics and Control Branch,
SeniorMember,AIAA single wind tunnel test, conducted from
November - December 1995, focused on the
Copyright © 2002bytheAmericanInstituteofAeronauticsand Reynolds number sensitivities of longitudinal
Astronautics,Inc. Nocopyright isassertedintheUnitedStates
aerodynamic characteristics at subsonic and
under Title 17, U. S. Code. The U. S. Government has a
transonic conditions for three distinct
royalty4ree license to exercise all rights under the copyright
claimedhereinforGovernmentalPurposes. Allotherrightsare configurations. The three major configurations
reservedbythecopyrightowner.
1
American Institute of Aeronautics and Astronautics
AIAA-2002-0418
tested were 1) a standard Fuselage / Wing (FW) EXPERIMENTAL APPROACH
configuration, 2) a Fuselage / Wing / Centerline Facility Description
Vertical Tail / Horizontal Tail (FWV1H) The NTF is a unique national facility (fig.
configuration, and 3) a Fuselage / Wing / 1) that enables testing of aircraft configurations
Trailing-Edge Extension / Twin Vertical Tail at conditions ranging from subsonic to low
(FWTV2) configuration. supersonic speeds at Reynolds numbers up to
full-scale flight values, depending on the aircraft
type and size. The facility (fig. 2) is a fan-driven,
TERMS, ABBREVIATIONS. & closed-circuit, continuous-flow, pressurized wind
ACRONYMS
tunnel capable of operating in either dry air at
warm temperatures or nitrogen from warm to
AF Axial force
cryogenic temperatures. The test section is 8.2
AR Aspect ratio ft by 8.2 ft in cross section and 25 ft in length.
BL Butt line
The test section floor and ceiling are slotted (6
CI95 95% confidence interval
percent open), and the sidewalls are solid.
c Local chord length Freestream turbulence is damped by four
CD Drag coefficient screens and a 15:1 contraction ratio from the
CL Lift coefficient settling chamber to the test section. Fan noise
CM Pitching-moment coefficient effects are minimized by an acoustic treatment
referenced to 0.42 mac both upstream and downstream of the fan. The
CP Pressure coefficient NTF is capable of an absolute pressure range
CR Root chord from 15 psi to 125 psi, a temperature range from
CT Tip chord -260°F to 130°F, a Mach number range from 0.2
E Modulus of elasticity to 1.2, and a maximum Reynolds number of
FW Fuselage / Wing configuration 146x106 per ft at Mach 1. Further facility details
FWV1H Fuselage / Wing / Centerline can be found in reference 5.
Vertical Tail / Horizontal Tail
configuration
Model Description
FWVT2 Fuselage / Wing / Trailing-Edge
The Pathfinder II configuration with the
Extension / Twin Vertical Tail
McDonnell Douglas wing is a generic model of a
configuration
general research configuration. The model,
LaRC Langley Research Center
shown in figure 3, has the capability for testing
LE Leading edge
numerous aerodynamic concepts including: two
M Mach number
segment leading-edge flaps, trailing-edge flaps
mac Mean aerodynamic chord
and ailerons, a trailing-edge extension (TEX)
MS Model station
with a trailing-edge flap, and a wing tiperon.
NF Normal force
Only the TEX, un-flapped configuration was
NTF National Transonic Facility
tested. The three configurations tested for this
PM Pitching moment
investigation are sketched in figure 4. Figure 5a
PT Total pressure
shows a planform drawing of the model with
Q Dynamic pressure
pertinent reference geometry denoted. The
Rn Reynolds number based on mac
model has a delta wing with an aspect ratio of
RM Rolling moment
1.946, a span of 20.802 inches, and a mean
SF Side force
aerodynamic chord of 13.434 inches. The wing
Sref Reference area
has a leading-edge sweep of 65 degrees with a
TT Total temperature trailing-edge sweep on the outboard panel of 35
TE Trailing edge degrees. The airfoil section is a NACA 65A004
TEX Trailing-edge extension at the root and a NACA 65A005 at the tip with a
YM Yawing moment linear thickness distribution from root to tip.
c_ Angle of attack
Figure 5b details the geometry of the respective
q Non-dimensional semispan
empennage components. The reference area for
station the model is 1.544 square feet.
2
American Institute of Aeronautics and Astronautics
AIAA-2002-0418
The model was designed and description of these measurements and
constructed specifically for testing in the subsequent calculations is given in reference 7.
cryogenic, pressurized conditions of the NTF
where dynamic pressures up to approximately Data Reduction and Corrections
2300 psf were required for this investigation. Information on the various
The model was mounted in the NTF test section instrumentation devices, the data acquisition
on a straight sting. The sting mounted to a stub and control computers, and the data reduction
sting which in turn mounted to the facility arc algorithms for the different measurement
sector resulting in a model angle of attack range systems is provided in reference 7. Standard
for the test from -2 to 18 degrees. Pertinent balance, angle-of-attack, and tunnel parameter
model geometry as compared to the NTF test corrections have been applied. Note that the
section geometry is shown in table 1. The use of unheated balances in the cryogenic
model was relatively small in comparison to environment requires additional attention
typical NTF transport or high-speed research towards temperature compensation. The
models. temperature compensation methods are
The model was instrumented with 43 designed to correct balance output due to
pressure ports on the wing upper surface and 18 thermal loads. Body cavity pressure corrections
on the wing lower surface. Additionally, there were applied based on the measurements
were 22 upper surface pressure ports along the described previously. The angle of attack was
fuselage centerline (aft of nose) and slightly off- corrected for flow angularity (upflow) by
center on the aft fuselage. Limited pressure measurement of both upright and inverted model
data is presented herein. force data for a given configuration and flow
condition. Wall and model support interference
Instrumentation effects have not been accounted for in the data;
Aerodynamic force and moment data wall effects are assumed minimal due to the
were obtained with an internal, unheated, six- model size relative to test section (see table 1).
component strain gauge balance. Design loads
for the NTF 104B balance were: NF=3400 Ibs, Test C,onditions
AF=300 Ibs, PM=10,000 in-fbs, RM=5000 in-fbs, The test program was set-up to evaluate
YM=5000 in-lbs, and SF=1000 Ibs. The quoted the effect of full-scale Reynolds numbers and
accuracy from the calibration of the balance was produce an aerodynamic database applicable to
less than or equal to 0.23% of full-scale load for thin-wing, fighter-type configurations. The NTF
each balance component. allows testing across a wide range of Reynolds
An internal, heated, single-axis, on- numbers from that available in conventional
board accelerometer package was used to wind tunnels to near flight conditions at subsonic
measure the model angle of attack. The and transonic Mach numbers. The Reynolds
accelerometer package had a quoted accuracy, numbers chosen for the test matched full-scale
under smooth operating tunnel conditions, of conditions at selected altitudes of 20,000,
±0.01 degrees (ref. 6). 30,000, and 40,000 feet, representative of
Model pressure measurements were operational fighter aircraft altitudes. Tests of the
obtained using two 48-port, 30-psid, onboard, Pathfinder II model spanned Reynolds numbers
heated, electronically scanned pressure (ESP) from 5 million to 60.9 million at Mach numbers
transducers with a quoted accuracy of ±0.2% of 0.6, 0.8, and 0.9. Representative test points for
full-scale pressure range. The body cavity Mach 0.9 are shown in figure 6. Data were
pressure was measured at two locations inside obtained at several total temperature conditions
the fuselage cavity with a heated, 2.5-psid ESP requiring both air and nitrogen mode operations.
module located inthe facility arc sector. Data were obtained over an angle-of-attack
The primary measured flow variables range from -2 to 18 degrees.
included both the total and static pressures and The initial configuration was tested to
the total temperature. Mach number, Reynolds acquire force and model pressure data
number, and dynamic pressure were calculated simultaneously. In an attempt to relieve any
from these measured parameters. A complete possible fouling or any thermally induced loads
from the pressure tubing bridging the balance,
3
American Institute of Aeronautics and Astronautics
AIAA-2002-0418
the pressure tubes were cut and removed after a Repeatability
set of pressure data was acquired on the initial The short-term repeatability of the force
Fuselage / Wing configuration. The and moment data from the Pathfinder II model
configuration was then re-tested for only force was analyzed. Repeat runs were not
and moment data. Hence, there is force and conducted for every configuration or for all Mach
moment data for all three configurations numbers, but were scattered throughout the test
available, but only pressure data available for program. The analysis for the available repeat
the Fuselage /Wing configuration. runs was conducted using the methodology as
described in reference 8. The analysis consists
Boundary-Layer Transition of statistical determination of the mean value of
A basic strategy used at the NTF the selected coefficients from repeated runs, a
includes testing at high Reynolds number curve fit of the data using a 2°dorder polynomial,
conditions with free transition. The high and a determination of the residual of the
Reynolds number test condition typically individual data points from the curve fit data.
corresponds to a design flight condition. To Confidence intervals are determined and are
anchor the NTF data to low Reynolds number defined as the bounds about an estimated mean
data obtained in a conventional wind tunnel, the that encompasses the true mean value with a
NTF model is usually tested at a matching low probability of 95 % confidence.
Reynolds number condition with the boundary- A repeatability residual analysis is
layer tripping (forced transition) strategy used in shown in figure 8 for the longitudinal coefficients
a conventional wind tunnel. However, for this at Mach 0.6 and a representative high Reynolds
investigation, no tripping of the boundary layer number condition. Table 2 is a summary of
was utilized for any test condition (Mach, short-term repeatability of the longitudinal
Reynolds number) since the test was focused on coefficients for the representative FW
the high Reynolds number data evaluation. The configuration. Repeatability was excellent for
impact of this decision will be discussed in the longitudinal coefficients for the angles of
succeeding sections. attack where attached flow is dominant, but
degrades somewhat for angles of attack greater
RESULTS & DISCUSSION than 4 degrees and for higher angles of attack
The purpose of this paper is to where separated flow dominates. In general,
document the Reynolds number sensitivities of there is good short-term repeatability for the
longitudinal aerodynamic characteristics for a longitudinal aerodynamic coefficients and is
generic fighter configuration at subsonic and comparable to the quoted balance accuracy.
transonic conditions. The three configurations Long-term repeatability or test-to-test
investigated were a baseline Fuselage / Wing repeatability data are not available to compare.
assembly (FW), a configuration with a centerline
vertical tail and a horizontal tail (FWV1H), and a Static Aeroelastic Effects
twin vertical tail configuration with a trailing-edge The investigation of aerodynamic effects
extension (FWTV2). Figure 7 shows on a model in a variable-pressure facility such
representative data for the three configurations as the NTF should take into account the
at a representative Mach number of 0.6 at a possible aeroelastic effects. These effects could
medium Reynolds number of 22 million and is mask the other aerodynamic effects, such as
provided to indicate the basic aerodynamic Reynolds number effects, which are being
characteristics of the configurations. The data, studied. Since the NTF is capable of controlling
as acquired at varying Reynolds numbers, Mach number, dynamic pressure, and
included the combined effects of aeroelastic temperature independently, the desired test plan
deformation and Reynolds number effects would be to test at comparable dynamic
because the conditions at which the data were pressures, or more specifically, comparable
acquired were at different dynamic pressure dynamic pressures divided by modulus of
levels in general; further discussion will address elasticity (Q/E) values. However, based on
the aeroelastic effects. limitations on the strength of the model material,
tunnel capability limitations, or nitrogen usage
rates, the test plan is normally compromised,
4
American Institute of Aeronautics and Astronautics
AIAA-2002-0418
and the data from the test, if required, are above 22 million. Upon closer examination, the
"normalized" to provide similar wing shape drag for the lower Reynolds numbers (5 million
comparisons. and 12 million) does not appear to be consistent
Data for this particular model were taken with a general decrease in drag with increasing
at two different dynamic pressure levels at the Reynolds number as seen from the expected
same Reynolds number (Rn=22 million). A theoretical fully turbulent drag result (anchored
representative example of dynamic pressure to the highest Reynolds number). The
effects is shown in figure 9 for the FW measured drag for this condition is considerably
configuration. Data are presented as a delta below the theoretical, fully turbulent drag. The
plot and compare the higher dynamic pressure likely cause for this result is the lack of fixed
run at the same Reynolds number as referenced boundary layer transition at the lower Reynolds
to the lower dynamic pressure level. number conditions. For these conditions, it is
Aeroelastic effects on the longitudinal expected that the boundary layer flow over the
aerodynamic coefficients were detectable, even wing will be a combination of both laminar and
for this low aspect ratio configuration, but were turbulent flow. In fact, there can be a significant
generally small. It was decided not to adjust region of laminar flow for these conditions. This
data for these effects in the analysis for this would likely explain the apparent drag deficit
report, though future analysis may include such measured at this condition. To verify these
adjustments. It is interesting to note, in figure 9, results, a repeat of the baseline configuration
the increase in lift at high dynamic pressure (and with the boundary layer tripped in the NTF would
Q/E) even at high lift conditions, implying a be needed. There is no observable effect on
leading-edge-up wing twist not typically seen either lift or pitching moment at this particular
with other models. Though not shown, this trend angle of attack as a function of Reynolds
was not consistent for all Mach numbers and number.
may be indicative of changing center-of- The general Reynolds number trends for
pressure characteristics relative to the elastic the longitudinal aerodynamic coefficients near
axis of the wing. Wing twist photogrammetry minimum drag ((x = 0.5 degrees) at Mach = 0.8
measurements under load would have added to and 0.9 (figures 11, 12) for the FW configuration
understanding of these effects, but the currently are consistent with the trends at Mach= 0.6.
operational measurement system (ref. 9) was The drag coefficients at the lowest Reynolds
not available during this test. numbers (5 and 12 million) again show a
decreased magnitude as compared to the fully
Reynolds Number Effects turbulent theory. The pitching moment and lift at
The effects of Reynolds numbers on this these Mach numbers are also insensitive across
generic fighter configuration were analyzed for the Reynolds number range. There appear to
the three configurations and three different Mach be no compressibility effects on the trends with
numbers tested. Three distinct angles of attack Reynolds number for characteristics of this
were investigated to evaluate these effects for configuration at the minimum drag condition.
different flow states. These representative Figures 13 and 14 show Reynolds
angles of attack were 1) _ = 0.5 degrees, number trends for the FW configuration at a
approximately minimum drag, 2) (x = 4.0 near design, primarily attached flow condition (_
degrees, a near design condition, and 3) c¢ = = 4.0 degrees) and a highly separated flow
12.0 degrees, a separated flow condition. The condition (cz= 12.0 degrees) at Mach 0.90. The
highlights of this analysis are detailed below with expected effect of decreasing drag with an
key representative figures. increase in Reynolds number was observed,
and, though theory is not shown, there are less
Fuse(cid:0)age / W/ng Conh'guration (FW) indications of significant laminar flow at the low
For the baseline FW configuration at a Reynolds numbers. Lift is insensitive to
Mach number of 0.6, near minimum drag (_ = Reynolds number at these flow conditions, like
0.5 degrees), there appears to be an initial at the minimum drag condition, but the pitching
insensitivity to Reynolds number below 22 moments show an increasing sensitivity to
million, as shown in figure 10, and a trend of Reynolds number changes. At an angle of
decreasing drag as Reynolds number increases attack of 4.0 degrees (near the design condition)
5
American Institute of Aeronautics and Astronautics
AIAA-2002-0418
at Mach 0.90, the pitching moment (figure 13) Fuselage / Wing / Centerline Vertical Tad/
shows an increase (nose-up) as the Reynolds Horizontal Tail Conh'guration (FWV IH)
number increases beyond 12 million. This The effects of Reynolds number on the
characteristic is also observed at an angle of FWV1H configuration at Mach 0.9 and 4
attack of 12.0 degrees (figure 14), but the degrees angle of attack are shown in figure 16.
insensitivity below 12 million is gone and the Drag and lift trends with Reynolds number for
rate of increased pitching moment with Reynolds this configuration are very similar to the FW
number is higher. Note that there is a potentially configuration, though absolute levels change
significant aeroelastic effect at this separated due to the addition of model components. With
flow condition at a constant Reynolds number of increasing Reynolds number, the drag shows a
22 million. At this Reynolds number and angle downward trend; this trend holds at other angles
of attack, drag decreases with increasing of attack and for Mach 0.6 and 0.8. The lift is
dynamic pressure with a corresponding insensitive to Reynolds number for this
decrease in lift. This would be representative of configuration for all Mach numbers at 4 degrees
a small wing twist deflection downward angle of attack.
producing a decreased local angle of attack. The pitching moment trends for the
The aeroelastic effect is also apparent in the FWVIH configuration are different than the FW
pitching moment where an increased nose-up configuration. Figure 16 shows a decreasing
moment accompanies the increase in dynamic (more nose-down) moment with increasing
pressure; this result is consistent with a lift Reynolds number below 22 million; above 22
decrease associated with a leading-edge down million, pitching moment is insensitive to
wing twist at the tip with constant body angle of changes in Reynolds number. This different
attack. behavior, compared to the FW configuration, is
Another way of understanding the flow likely due to the addition of the tail components
characteristics as Reynolds number changes is and flow field interactions between the wing and
with the analysis of the model surface pressure tail. This configuration warrants further study to
data. Although there were limited pressure data understand the details relevant to this behavior.
acquired during the test, some interesting results
were obtained. For example, leading-edge Fuselage /Wing / Trailing-EdgeExtension/
pressures for one spanwise station (TI= 0.22) Twin VerticalTailConhguration (FHITV2)
are shown in figure 15 for the FW configuration. The effectsof Reynolds number on the
This figure highlights pressures for three FWTV2 configuration at Mach 0.9 and 4 degrees
Reynolds numbers as a function of angle of angle of attack are shown in figure 17. Drag and
attack at a Mach number of 0.9. Two trends can lift trends with Reynolds number for this
be seen. First, the leading edge pressures are configuration are very similar to the FW and
becoming increasingly more negative until the FWVlH configurations, though absolute levels
angle of attack reaches about 4 degrees. This is change due to the different empennage
consistent with attached flow below that angle of components. With increasing Reynolds number,
attack. Second, the effect of increasing the drag shows a downward trend; this trend
Reynolds number is to reduce the level of holds at other angles of attack and for Mach 0.6
suction for angles of attack greater than 4 and 0.8. The lift is insensitive to Reynolds
degrees. This reduction is consistent with the number for this configuration for all Mach
expectation that higher Reynolds numbers will numbers at 4 degrees angle of attack.
delay separation over the leading edge. The pitching moment trends for the
Delayed separation will result in less vorticity FWTV2 configuration are similar to the FWV1H
being generated and, therefore, weaker suction configuration, and thus different than the FW
in the separated region. This observation, configuration. Figure 17 shows a decreasing
however, does not help to explain the pitch-up (more nose-down) moment with increasing
character with Reynolds number noted above in Reynolds number below 22 million; above 22
figures 13 and 14. This discrepancy will be million, pitching moment is insensitive to
addressed in future studies. changes in Reynolds number at this angle of
attack. As with the FWV1H configuration, this
different behavior, compared to the FW
6
American Institute of Aeronautics and Astronautics
AIAA-2002-0418
configuration, is likely due to the addition of the CONCLUDING REMARKS
tail components and flow field interactions A wind tunnel test with a generic fighter-
between the wing and tail. This configuration type model was executed in the National
also warrants further study to understand the Transonic Facility at NASA LaRC across a wide
details relevant to this behavior. range of Reynolds numbers from that available
The neutral point characteristics of the in conventional wind tunnels to flight conditions
three configurations at Mach 0.9 are shown in at subsonic and transonic Mach numbers.
figure 18 as a function of Reynolds number. Results were presented which focus on the
The neutral point is defined as the location Reynolds sensitivities of longitudinal
where the pitching moment coefficient is characteristics at Mach 0.6, 0.8, and 0.90 for
independent of the angle of attack. It is also three unique configurations. General
representative of where the center of pressure conclusions are summarized as follows:
(center of lift) is located relative to the mean 1. Static aeroelastic effects on the longitudinal
aerodynamic chord. The pitching moment data aerodynamic coefficients were detectable,
were referenced to 0.416 mac. Figure 18 but were generally small. The effects were
shows, for the three configurations, the larger atthe higher angles of attack.
respective movement of the center of pressure 2. Drag trends for the lowest Reynolds
for angles of attack previously analyzed (0.5, 4, numbers (5 and 12 million) tested were
and 12 degrees). At a representative Mach differed from theoretical fully turbulent skin
number of 0.9, the baseline FW configuration friction estimates. The likely cause for this
shows a forward movement of the center of result is significant regions of laminar flow
pressure (about 4 %) for the two attached flow due to the lack of fixed boundary layer
conditions (o_=0.5 and (z=4.0 degrees). This transition at the lower Reynolds number
movement of the neutral point indicates an conditions.
increase in aft separation and decrease in aft 3. Lift was generally insensitive to Reynolds
loading with an increase in Reynolds number. number for the configurations and conditions
For the separated flow condition (c_=12 tested.
degrees), the center of pressure moves forward 4. Reynolds number effects are larger at
approximately 15 %. In general, for separated angles of attack where separated flow
flow conditions, the neutral point is not expected dominates, particularly the pitching moment
to maintain its position. However, this Reynolds characteristics. The presence of tail
number effect, in addition to the effect of the components altered the pitching moments
angle of attack, could have significant trends with Reynolds number.
implications on the stability and control power
required for the vehicle. The FWTV2 ACKNOWLEDGEMENTS
configuration shows a lesser, but still significant, The authors would like to thank the
effect at the higher angle of attack; center of Boeing Company, formerly McDonnell Douglas,
pressure movement at the attached flow for their support in this wind tunnel investigation.
conditions is similar to the FW. The FWVIH We would like to specifically thank Mr. Wayne L.
configuration significantly reduces the Ely (retired) of McDonnell Douglas for his
movement of the neutral point with Reynolds important contributions before, during, and after
number for the separated flow condition, but this investigation. Without his effort, this test
shows similar attached flow characteristics. would not have been accomplished. Also, we
would like to acknowledge the representatives of
the National Transonic Facility for their support
during the testing and subsequent data
processing.
7
American Institute of Aeronautics and Astronautics
AIAA-2002-0418
BEEEBEBC_
ref. area / NTF cross sectional area 0.023
1. Campbell, J.F.: "The National Transonic
model span / NTF width 0.211
Facility - A Research Perspective," AIAA
Paper 84-2150, August 1984. solid blockage ratio, (_= 0 deg 0.0039
2. Wahls, R.A.: "The National Transonic
Table 1. Model size relative to the NTF test section.
Facility: A Research Retrospective," AIAA
Paper 01-0754, January 2001.
3. Luckring, J.M.: "An Overview of National Math Rn, 106 q, psf _CD _CL _CM
Transonic Facility Investigations for High
0.6 5 59O .00004 .0003 .00006
Performance Military Aerodynamics," AIAA
Paper 01-0906, January 2001. 0.6 56 1608 .00020 .0020 .00010
4. Ely, W.L.: "Summary Report for the 0.9 60.5 2309 .00015 .0015 .00030
Pathfinder II Full-Scale Reynolds Number
National Transonic Facilty Wind Tunnel Test Table 2. Repeatability data for FW configuration (95%
- Test 77," Report Number MDA 96P0049, confidence interval evaluated over entire alpha range).
October 1996.
5. Fuller, D.E.: "Guide for Users of the National
Transonic Facility," NASA TM-83124, 1981.
6. Finley, T.D. and Tcheng, P.: "Model Attitude
Measurements at NASA Langley Research
Center," AIAA Paper 92-0763, 1992.
7. Foster, J.M. and Adcock, J.B.: "User's Guide
for the National Transonic Facility Research
Data System," NASA TM-110242, April
1996.
8. Wahls, R.A., Adcock, J.B., Witkowski, D.P.,
and Wright, F.L..: "A Longitudinal
Aerodynamic Data Repeatability Study for a
Commercial Transport Model in the National
Transonic Facility," NASA TP-3522, August
1995.
9. Burner, A.W., Wahls, R.A., and Goad, W.K.:
"Wing Twist Measurements at the National
Transonic Facility," NASA TM-110229,
February 1996.
8
American Institute of Aeronautics and Astronautics