Table Of ContentAIAA-98-2725
IMPACT OF FUSELAGE CROSS SECTION ON THE
STABILITY OF A GENERIC FIGHTER
Robert M. Hall?
NASA Langley Research Center
Hampt on, Virginia
ABSTRACT
Many traditional data bases, which involved smooth-sided forebodies, are no longer relevant for designing
advanced aircraft. The current work provides data on the impact of chined-shaped fuselage cross section on the
stability of a generic fighter configuration. Two dfferent chined-shaped fuselages were tested upright and inverted.
It was found that a fuselage with a 30" included chine angle resulted in significantly higher values of than a
fuselage with a 100" included chine angle. This difference was attributed to the more beneficial vortical interaction
between the stronger forebody vortices coming off of the sharper chine edges and the wing vortices. The longitudinal
stability of the configuration with the sharper chine angle was also better because, based on pressures and flow
visualization, the vortex burst over the wing was delayed until significantly higher values of a. Unstable rolling
moment derivatives were also delayed to higher values of a for the sharper chine angle cross section. Furthermore,
it was found that drectional stability of both of the upright configurations, which had larger lofts in cross section
above the chine lines than below the chine lines. was better than for the inverted configurations.
INTRODUCTION longitudinal and lateral/directional stability of a
generic fighter model. The overall study involved a
Current advanced fighter designs usually number of wind tunnel entries at subsonic and
include a fuselage with chine-shaped cross section to transonic speeds. A total of 4 aferent fuselage cross
minimize observables. The presence of the chines sections were evaluated. Two of these four fuselages
results in a fixed location of flow separation and were tested inverted. The 4 fuselages were tested
generally stronger forebody vortices being shed than with either a centerline vertical tail, twin vertical
would be the case for a configuration with a smooth- tails, or no tails. All the fuselages had the same
sided forebody. These stronger vortices can be useful planform and utilized the same cropped delta wing
in that they can synergistically augment the lift over with a 60" leading-edge sweep. The wing itself was
the main wing; however, these chine vortices can also fitted with leading-edge flaps. The test facility, the
lead to pitch up and nonlinearities in lateral stability Langley 7- by 10-Foot High Speed Tunnel (HST),
due to windward vortex burst at sideslip. permitted testing to an angle of attack of up to 50"
Consequently, vortices from chined forebodes can and with sideslip values as large as ? 15". Mach
have a significant impact on both longitudinal and number in th~fsa cility was, for the purposes of th~s
1ateraL'directional stability. test, limited to 0.4. For the present paper, data from
the 7- by 10-Foot HST will be reported for two of the
The present paper describes a portion of a fuselage cross sections, each tested upright and
joint NASA and U. S. Air Force study undertaken at inverted. The data for the configurations with no
the Langley Research Center to investigate the impact leadng-edge flaps deflections and no vertical tails
of chine-shaped fuselage cross section on the will be presented.
The data obtained during this experimental
?Senior Aerospace Engineer, Configuration Aerodynamics
Branch, Associate Fellow, AIAA program have been extensively used for validation of
Euler and Navier-Stokes numerical codes. Example
Copyright 0 1998 by the American Institute of Aeronautics and publications utilizing these data for code validation
Astronautics, Inc. No copyright is asserted in the United States
under Title 17, U.S. Code. The U.S. Government has a royalty- include references 1 to 3. While initial experimental
free license to exercise all rights under the copyright claimed results with one configuration were presented in
herein for government purposes. All other rights are reserved by
the copyright owner. references 4 and 5, this report will examine the
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impact of changing the shape of the fuselage cross DESCRIPTION OF WIND TUNNEL AND
section on the vehicle stability. MODELS
NOMENCLATURE Wind Tunnel Facilitv
The longitudinal data are referred to the stability-axis The NASA Langley Research Center 7- by
system and the lateral-directional data are referred to 10-Foot HST is a continuous-flow, solid-wall,
the body-axis system. The data are normalized by subsonic/transonic atmospheric wind tunnel, which is
the usual quantities, such as planform area, span of described in more detail in reference 6. The model
the wing, and the wing mean aerodynamic chord. was supported on a sting system with roll
The moment reference center was located at the 0.25 positioning capability. A combination of pitch and
mean aerodynamic chord location, which was at roll was used to generate the desired combinations of
model fuselage station 20.36 inches. a and B.
The tunnel was equipped with a laser light
reference wing span, 19.20 in sheet capability that was used during the present
body-axis axial force coefficient, program to visualize the vortex flow development
Axial Force over the model. This unpublished video record was
invaluable in understanding the relationship between
9,s
the force and moment data, the pressure data, and the
Drag
stability-axis drag coefficient, - complex vortical flow above the model.
9,s
Lift Wind Tunnel Model
stability-axis lift coefficient, -
9,s
maximum value of CL The wind tunnel model is shown in
body-axis rolling-moment coefficient, schematic form in figure 1. The model has a span of
Rolling moment 29.20 inches and a length of 32.48 inches. The wing
has a 60" cropped-delta planform incorporating a
9 s
derivative of C1 with sideslip, B, per deg NACA 65-005 airfoil mo&fied with a double-arc
section forward of the maximum thdness and has
body-axis pitching-moment coefficient,
sharp leading edges. The fdl-span leading edge flaps
Pitching moment
are divided into 3 segments of equal span, where only
9,s;
the 2 inner segments were built to be deflected. The
body-axis yawing-moment coefficient, model system included options for a centerline
Yawing moment vertical tail and twin, wing-mounted vertical tails. A
9,Sb total of 276 static pressure orifices were grouped into
derivative of Cnw ith sideslip, B, per deg six spanwise rows with three on the forebody and
three on the wing. The common planform for the two
Side force
body-axis side force coefficient, configurations to be reported, which were tested
9,s
upright and inverted, is shown in figure 2. The
center of gravity model was mounted in the tunnel with a 6-
wing mean aerodynamic chord, 12.97 in. component force and moment balance. For the present
free-stream Mach number report, data for the tails-off configuration with no
National Aeronautics and Space lea&ng edge flap deflection will be presented.
Administration
free-stream dynamic pressure, lbh' The fuselage cross-sections are highlighted
Reynolds number based on in figure 3. The two fuselages with total included
reference wing area, 208.19 in' chine angles of 30" and 100" were run both upright
angle of attack, deg and inverted. Since the wing sections are symmetric
angle of sideslip, deg top to bottom, running the model inverted highlights
the effect of inverting the forebody cross sections,
which are not symmetric top to bottom. Details of
the fuselage cross sections are included in the
following table, where the loft values are the &stance
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either to the top or bottom of the fuselage from the Figure 6 is the corresponding data for the
plane of the chines and is normalized by the local upright and inverted 100" chine angle configuration.
fuselage half width. A photograph of the model with For th~cso nfiguration, the weaker chine vortices
the 30" chine angle installed in the wind tunnel is associated with the larger chine angle do not
shown in figure 4. synergistically interact with the wing vortices to the
extent that the chine vortices did with fuselage 1.
Fuselage Total Upper Lower Loft Loft This lack of interaction is evident in the earlier wing
chine chine chine above below vortex burst location both in the flow visualization
angle angle angle chine chine (not shown) and in the breaks in the lift curve slope
1 30" 20" 10" 0.75 0.5 at lower values of a for this second fuselage. For th~s
configuration, CL,maisx only 1.4 for the upright
2 100" 60" 40" 0.75 0.5
configuration and 1.5 for the inverted configuration.
1. Inverted 30" 10" 20" 0.5 0.75
II I I I I II The respective angles at which vortex bursting
2. Inverted 100" 40" 60" 0.5 10.75
occurred are approximately 25" and 27.5", based on
I ' I I I I I I
the laser light sheet data, which is why the value of
CL,maisx lower for the upright configuration. The
RESULTS AND DISCUSSION reason why the inverted configuration had a larger
value of CL,mathxa n did the upright configuration
The results will be shown for M, = 0.40. while the opposite is true for 30" chine configuration
The figures will present the a-polar and fl-polar force is not known.
and moment data for the fuselages with the 30"
(fuselage 1) and 100" (fuselage 2) total included chine Both the upright and inverted 100" chine
angles, upright and inverted, and then calculated fl- configurations have a more unstable pitching moment
derivative data to illustrate the lateral/directional character even at the lower values of CLt han do the
stability of the configurations. The fuselage with the upright and inverted configurations with the 30" chine
smaller included chine angle, 30", upright and angle. An interesting feature with the two 100" chine
inverted, is expected to generate stronger chine configurations is that there are some pronounced
vortices over the forebody than the fuselage with the lobes in the plots of Cy, G, and C, versus a even
100" chine angle, see reference 7. though the configuration was nominally at fl = 0".
Based on flow visualization and wing pressure data,
Figure 5 shows the a-polar data at fl = 0" for these lobes correspond to the burst of the wing vortex
the upright and inverted 30" chine angle on one side of the configuration. The beginning of
configuration. The value of CL,mafoxr the upright this asymmetric vortex process is most evident in the
configuration is approximately 1.9 and occurs at a plot of C1 versus a, where it is clear that the vortex
value of a = 38". The corresponding value of CL,max asymmetry is occurring concurrently with the breaks
for the inverted configuration is about 1.8 and occurs in the lift curve slope.
at a = 35". There is no vortex bursting for either the
forebody or wing vortices up to CL,mabxa, sed on the The character of both the 30" and 100" chine
unpublished laser light sheet data. In the plots of CL angle configurations in sideslip is illustrated in
versus C,, the curves are very similar, but there is an figures 7 and 8. The upright and inverted data for the
offset in C, that may be the result of the change in small chine angle of 30" are shown in figure 7.
effective fuselage camber when the fuselage is Significant hysteresis occurs for these configurations
inverted. As shown in the side view of figure 2, with in sideslip. In fact, in the figure for C1 versus fl at
the higher upper loft of the upright fuselage, a mean a = 23", it can be seen that the data for the upright
fuselage line in side view would have an effective configuration (circles) are Merent for negative and
nose droop. Conversely, the inverted fuselage would positive values of fl. The aference between the data
have an effective nose-up appearance. Consequently, is due to the direction in which fl was varied. The
one would expect to see a positive pitch increment values of fl were varied from large negative values to
when changing from the upright fuselage to inverted large positive values in one continuous sweep.
fuselage configuration, as indicated. The non-zero When the data were repeated later by varying fl so
values of Cy, G, and C, are indicative of asymmetries that the configuration was always going from zero
right to left either in the model geometry or sideslip to larger magnitudes of sideslip, it was found
installation or are indcative of asymmetric flow that the character for C1 versus fl without hysteresis
Merences. for a = 23" was the one on the positive fl side of the
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AI AA-98-2725
curve shown in figure 7. The problem with the calculated using values in the neighborhood of fl = 0.
negative values of fl was that the windward vortex That is, in any differencing using the data about
system, composed of the chine and wing vortices, fl = 0", the derivative obtained will vary dramatically
was already burst at the higher negative values of fl depending on whether G = +.04, 0, or -.04.
and did not reform until fl > -2". It is the unburst
character of the windward chine and wing vortices The correspondmg values of C, are shown to
that give the large negative values of G near fl = 4" the right in figure 7 and suggest that the inverted
for both the upright and inverted configurations. configuration is less directionally stable (stable is
(These upright data at a = 23" are the only fl-polar defined as Cnp positive) for values of IflI > 3" for all
data which were sampled by a continuous sweep of fl
three values of a. This would be expected because
from negative to positive. All other fl-polars shown
the inverted configuration has less side area above the
were sampled by starting at fl = 0" and then chine line on which the windward chine vortex can
proceeding to either negative or positive values.)
act. It is the windward chine vortex which, in
sideslip, reduces the pressure on the windward
The trends of the G versus fl plots in figure
fuselage side (confirmed by unpublished pressure
7 can be readily explained with the assistance of the
measurements not shown) and is responsible for the
laser light sheet data that was simultaneously
restoring moment in C,.
acquired. For both a = 23" and 30", there are clear
regions surrounding fl = 0" that illustrate lateral The correspondmg Cl data for the fuselage
stability (negative Clp).F or a = 23", th~sre gion is
with the 100" chine angle is shown in figure 8. For
-4" < fl < 4". For a = 30", this region is the upright configuration at a = 23", there is no
-2" < fl < 2". It is expected that this region may also region of local stability around fl = 0". The inverted
exist for a = 35", but that it is over such a small configuration does better and exhibits a stable region
range of fl that it is not evident in the figure. for IflI < 2". Neither the upright or inverted
Physically, these regions correspond to the range of fl
configurations for the 100" chine angle fuselage have
where both the windward wing and chine vortices are local regions of stability around fl = 0" at a = 30" or
not burst. At larger magnitudes of fl, the windward 35". Multiple values in G near fl = 0" are evident at
chine and wing vortices, which are interacting over
a = 30" and 35" for the upright configuration. While
the wing, burst and the windward wing loses lift.
there are not multiple values of G for the inverted
This results in the large jumps in Cl. Note also that configuration, the fact that the G values for fl = 0"
while there are laterally stable regions near fl = 0" for are not near zero for a = 30" and 35" suggests the
a = 23" and 30", the behavior of Cl versus fl results possibility of multiple values if more data had been
in regions of lateral stability about non-zero values of
taken or if the model had been reinstalled in the
fl at a = 30" and a = 35". For a = 30", the tunnel.
configuration would tend to "fly" near IflI = 5"
(Cl = 0) if it exceeded values of IflI greater than 2". As was the case for the 30" chine
At a = 35', it may be that any perturbation in configuration, the upright 100" chine configuration
sideslip would tend to take the configuration out to demonstrates higher restoring moments of C, over the
IflI = 4". At this value of a, the configuration may range of fl examined. In general, both upright and
not practically be flyable at fl = 0" with such a sharp inverted configurations are stable for IflI < 10" with
discontinuity in G around fl = 0". With the some nonlinearities near fl = 0".
sharpness of the chine edge in either the upright or
inverted orientation, the differences in the strength of A more drect look at lateral and directional
the chine vorticity may be relatively small and do not stability is given in figures 9, 10, and 11. The
appear to result in significant change to the values of Clp and Cnp were calculated by Merencing
interactions between the chine and wing vortices. a-polars taken at fl = -2" and 0". The value of
fl = -2" was chosen to utilize the large excursions in
In figure 7, hysteresis is also responsible for G that occur for all four configurations near this angle
the multiple values of G near fl = 0" at a = 35". of sideslip. The comparisons in figure 9 are for the
These multiple values at a = 35" could be expected 30" chine angle configurations and are consistent with
to result in multiple values in any fl-derivatives of G the observations already made. That is, the lateral
stability of the upright and inverted configurations are
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very similar. The break in the curves at a = 30" interaction between the stronger forebody vortices
correlates with the onset of bursting of the wing and coming off of the sharper chine edges and the wing
chine vortex system over the windward wing for vortices. The longitudinal stability of the sharper
IflI 2 2". In terms of directional stability, the higher chine angle was also better because, based on
levels of directional stability of the upright unpublished pressures and flow visualization, the
configuration indicated in figure 7 are again indicated vortex burst over the wing was delayed until
in th~spl ot of Cnp versus a. significantly higher values of a. Rolling moment
derivatives were also much more stable for a 5 30"
The comparable derivative data for the 100"
for the sharper chine angle cross section. Because of a
chine angle configuration is shown in figure 10. For
lack of repeatability, it was inconclusive whether the
th~fsu selage cross section, differences occur in Clp
upright or inverted fuselage orientations gave higher
between the upright and inverted configurations that values of lateral stability for either fuselage.
are quite pronounced. First, the upright configuration Nevertheless, it was found that directional stability of
has stable, negative values of Clp over the a-range, both upright configurations, which had larger lofts in
while the inverted configuration has an unstable cross section above the chine lines than below the
region for 25" < a < 3 1". However, as was the case chine lines, was better than for the inverted
for the 30" chine configuration at a = 35", the values configurations. The increased directional stability
of Clpf or the upright 100" configuration could be was attributed to the larger side area above the chine
on which the windward chine vortex could act.
multivalued at a = 23 " because of the multiple However, the lack of repeatability and the severe
values of CI near fl = 0", seen in figure 8. nonlinearities in lateral stability for all four
configurations would be a sipficant challenge in
This multivalued nature of the upright 100"
developing flight control systems.
configuration is confirmed in figure 11, which repeats
the data of figure 10 and adds, for the upright
REFERENCES
configuration, a second derivative calculation based
on another set of a-polars at fl = 0" and -2". This 1. Treiber, David A,; and Muilenburg, Dennis A,:
lack of repeatability makes conclusions difficult
Euler Technology Assessment for Preliminary
regardmg the worth of inverting this 100" chine
Aircraft Design Employing OVERFLOW Code With
configuration with regard to lateral stability.
Multiblock Structured-Grid Method. NASA
Contractor Report 4651, March 1995.
As was the case for the 30" configurations,
the directional stability of the 100" configuration was
2. Finley, Dennis B.; and Karman, Steve L., Jr.:
slightly better for the upright configuration. Again,
Euler Technology Assessment for Preliminary
this is assumed to be due to the greater area above the
Aircraft Design--Compressibility Predictions by
chine on which the chine vortex can act.
Employing the Cartesian Unstructured Grid
SPLITFLOW Code. NASA Contractor Report
In comparing the lateral stability of the two
4710, March 1996.
fuselages, it appears that the lateral stability of the
100" configuration is not as good as that of the 30"
3. Kinard, Tom A,; and Raj, Pradeep: Euler
configmationfor a 5 30". However, for Technology Assessment for Preliminary Aircraft
32.5" 5 a 5 37.5", the lateral stability of the 100" Design--Compressibility Predictions by Employing
configuration appears to be better.
the Unstructured Grid USM3D Code. NASA
Contractor Report 471 1, March 1996.
CONCLUSIONS
4. Rao, Dhanvada M.; and Bhat, M. K.: Subsonic
An experimental study investigated the
Investigations of Vortex Interaction Control for
impact of forebody cross section geometry on the
Enhanced High-Alpha Aerodynamics of a Chine
stability of a generic fighter configuration. It was
ForebodyDelta Wing Configuration, NASA
found that a chined shaped fuselage with the 30"
Contractor Report 189641, June 1992.
included chine angle (the sharper chine) resulted in
sipficantly higher values of than a fuselage
5. Rao, Dhanvada M.; and Bhat, M. K.: High-
with a 100" included chine angle. This difference
Alpha Vortex Decoupling Investigations on a Chine
was attributed to the more beneficial vortical
ForebodyDelta Wing Configuration at Transonic
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Mach Numbers. NASA Contractor Report 189642,
June 1992.
6. Penaranda, Frank E.; and Freda, M. Shannon:
Aeronautical Facilities Catalogue. Volume I-- Wind
Tunnels. NASA RP-1132, 1985.
7. Nutter, J.: Leading-Edge Separation from a
Thick, Conical, Slender Wing at Small Angles of
Incidence. Journal of Engineering Mathematics, Vol.
15, April 1981, pp. 103-117.
Figure 3. Schematic of forebody cross sections in the
Figure 1. Schematic of overall model system.
upright orientation.
Figure 2. Planform and side view of model system.
Figure 4. Installation of 30" chine angle fuselage
configuration in the Langley 7- by 10-Foot HST.
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'
Facility Chine Orientation M, R?/IO~
7x10-FootHST 30. Upright 0.401 2.61
n--- - - - - ~ 7x10-Foot HST 30. Inverted 0.401 2.84
2.0 2.0
1.5 1.5
1.0 1.0
CL CL
.5 .5
0 0
-.5 -.5
-10 0 10 20 30 40 50 .4 .3 .2 .1 0 -.l
cm
a, deg
2.0
1.5
1.0
CL
.5 -.02 ~ ~ ~ -
-
~ ~ ~
-.04
0
-.5
.4 .X 1.2 1.6 2.0 2.4
CD
.04
.02
c1 0
-.02
-.04
-10 0 10 20 30 40 50
a, deg
Figure 5. 30" chine fuselage. Force and moment data for upright and inverted configurations. fl= 0".
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Facility Chine Orientation M, R~/l0~
7x10-FootHST 100. Upright 0.402 2.67
o-------- 7x10-Foot HST 100. Inverted 0.401 2.69
2.0 2.0
1.5 1.5
1.0 1.0
CL CL
.5 .5
0 0
-.5 .5
-10 0 10 20 30 40 50 .4 .3 .2 .1 0 -.1
c m
a, deg
2.0 .04
1.5 .02
1.0 CY 0
CL
.5 -.02
-.04
0
-10 0 10 20 30 40 50
a, deg
-.5
0 .4 .X 1.2 1.6 2.0 2.4
CD
.04 .04
.02 .02
c1 0 cn 0
-.02 -.02
-.04 -.04
-10 0 10 20 30 40 50 -10 0 10 20 30 40 50
a, deg a, deg
Figure 6. 100" chine fuselage. Force and moment data for upright and inverted configurations. 0 = 0".
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Facility Chine Orientation M, R~/l0~
~
0 7x10-FootHST 30. Upright 0.401 2.68
0 7x10-FootHST 30. Inverted 0.401 2.84
~
.05 .05
.04 .04
.03 .03
.02 .02
.01 .01
c1 0 Cn 0
-.01 -.01
-.02 -.02
-.03 -.03
-.04 -.04
-.05 -.05
-15 -10 -5 0 5 10 15 -15 -10 -5 0 5 10 15
B. deg B>d eg
.05 .05
.04 .04
.03 .03
.02 .02
.01 .01
c1 0 Cn 0
-.01 -.01
-.02 -.02
-.03 -.03
-.04 -.04
-.05 -.05
-15 -10 -5 0 5 10 15 -15 -10 -5 0 5 10 15
B. deg B>d eg
.05 .05
.04 .04
.03 .03
.02 .02
.01 .01
c1 0 Cn 0
-.01 -.01
-.02 -.02
-.03 -.03
-.04 -.04
-.05 -.05
-15 -10 -5 0 5 10 15 -15 -10 -5 0 5 10 15
B. deg B>d eg
Figure 7. 30" chine fuselage. Rolling and yawing moment data for upright and inverted configurations.
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Facility Chine Orientation M, F.~/l0~
0 7x10-FootHST 100 Upright 0.402 2.65
0 7x10-FootHST 100 Inverted 0.401 2.61
~
.05 .05
.04 .04
.03 .03
.02 .02
.01 .01
c1 0 Cn 0
-.01 -.01
-.02 -.02
-.03 -.03
-.04 -.04
-.05 -.05
-15 -10 -5 0 5 10 15 -15 -10 -5 0 5 10 15
P>d eg P>d eg
.05 .05
.04 .04
.03 .03
.02 .02
.01 .01
c1 0 Cn 0
-.01 -.01
-.02 -.02
-.03 -.03
-.04 -.04
-.05 -.05
-15 -10 -5 0 5 10 15 -15 -10 -5 0 5 10 15
P>d eg P>d eg
.05 .05
.04 .04
.03 .03
.02 .02
.01 .01
cll
c1 0 0
-.01 -.01
-.02 -.02
-.03 -.03
-.04 -.04
-.05 -.05
-15 -10 -5 0 5 10 15 -15 -10 -5 0 5 10 15
P>d eg P>d eg
Figure 8. 100" chine fuselage. Rolling and yawing moment data for upright and inverted configurations.
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