AIAA 2008-1422
46th AIAA Aerospace Sciences Meeting and Exhibit
7 - 10 January 2008, Reno, Nevada
Aerodynamic Evaluation and Optimization of the Houck
Joined Wing Aircraft
C1C Brittany Oligney1 and C1C Margaret Frash1
Department of Aeronautics, US Air Force Academy, CO 80841
Dr. Thomas R. Yechout2 (Research Advisor)
Department of Aeronautics, US Air Force Academy, CO 80841
Abstract
Air Force Research Laboratories (AFRL) tasked the United States Air Force Academy (USAFA)
Department of Aeronautics with investigating the aerodynamic characteristics of the Houck Joined Wing
Concept Aircraft in their search for a long range, long endurance Unmanned Aerial Vehicle (UAV). Because
lift induced drag entails a large penalty on aircraft performance, the Houck aircraft connects the upper and
lower wings, removing the wingtips in an attempt to eliminate these effects. The first version of the design
was evaluated at USAFA in Fall 2006. Aerodynamic performance and static stability characteristics were
analyzed and compared to a conventional mono-wing NASA Orbital Space Plane model. Testing showed that
the Houck Joined Wing had a somewhat shallow drag polar, indicating a possible reduction of drag due to
lift. However, the Joined Wing had a 7 times higher value for minimum drag coefficient when compared to
the Space Plane. The Joined Wing reached static longitudinal stability at angles of attack above 9° but was
unable to trim for the configuration evaluated. In addition, it was unstable in roll at low angles of attack.
Roll stability improved with increasing angle of attack, becoming the most stable at 15°. The Joined Wing
was unstable in yaw for
all angles of attack except at 0° angle of attack at Mach numbers of 0.2 and 0.3. Due
3
CL 2
to its comparable
C ratio to the Space Plane, the Houck aircraft configuration may have potential as a
D
long endurance UAV with a reciprocating engine if operated at high lift coefficients. No potential was found
for improving range or endurance for jet powered flight or range for reciprocating engine power.
The Joined Wing’s initial design configuration was matched as closely as possible to Mr. Houck’s original
design. However, using general engineering rules of thumb and a University of Missouri study on the biplane
design, adjustments were made in an attempt to optimize the Joined Wing’s performance. Changes
considered included a negative decalage angle, a taper ratio less than one, increased gap, decreased wing
sweep, and / or decreased stagger. Making all these changes simultaneously would not allow the effects of
each to be seen; consequently, gap and decalage angle were chosen to be varied on a modified Joined Wing
model. The increased gap (4.75 in) with -1.5° decalage angle proved to be the optimum design configuration
tested, producing higher lift coefficients and a more shallow drag polar. None of the Joined Wing
configurations evaluated showed significant potential for performance improvement over a monoplane. In
fact, the test model produced the best results with the top wing removed altogether.
Nomenclature
CL =
CD =
Cm =
Cm =
1
2
Lift coefficient, unitless
Drag coefficient, unitless
Pitching moment coefficient, unitless
Longitudinal static stability derivative, per deg
Cl =
Cl =
Lateral static stability derivative, per deg
Cn =
Yawing moment coefficient, unitless
Rolling moment coefficient, unitless
=
Sideslip angle, deg
q=
Dynamic pressure, psi
S=
Planform area, sq in
N1 , N 2 =Normal force, lb
Ax = Axial force, lb
Y1 , Y2 = Side force, lb
=
P0 =
Ratio of specific heats, unitless
Total pressure, psi
Undergraduate Student, Department of Aeronautics, USAFA CO 80841, AIAA Student Member
Professor, Department of Aeronautics, USAFA CO 80841, AIAA Associate Fellow
1
This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.
Cn = Directional static stability derivative, per deg
L = Lift, lb
D = Drag, lb
Mp
My
Mr
=
=
=
Pitching moment, in lb
Yawing moment, in lb
Rolling Moment, in lb
L
=
Lift-to-drag ratio, unitless
D
=
Angle of attack, deg
P = Change in pressure, psi
M = Moment, in lb
c = Chord length, in
U
=
Uncertainty, %
Bx = Bias error, %
Px = Precision error, %
I.
Introduction
Drag forces acting on a subsonic aircraft can be divided into three categories: induced, pressure, and skin friction
drag. Induced drag degrades the aerodynamic performance of the aircraft especially in the low subsonic regime,
which is the operating range of most UAVs. The Joined Wing Design attempts to eliminate the induced drag caused
by wingtip vortices. This is accomplished using a basic biplane-type design and connecting the upper and lower
wings at the tips with flow guide vanes, thus eliminating the wingtips. However, a tradeoff results from the
elimination of induced drag because of the increase in skin friction drag due to added surface area of the flow guide
vanes. The Air Force Research Laboratory (AFRL) tasked the United States Air Force Academy (USAFA) with
evaluating the aerodynamic performance and static stability characteristics of the Joined Wing Design to determine
the potential of this design concept for long range, long endurance UAVs. AFRL / VAAA (Air Vehicles
Directorate) provided USAFA with a computer aided design (CAD) file of the design, which was used to fabricate a
wind tunnel model. Testing was accomplished in the USAFA Aero Lab Subsonic Wind Tunnel. Tunnel Vision, a
program developed at USAFA that converts force balance measurements into aerodynamic coefficients, was used
for data reduction. For the3 longitudinal analysis,
lift, drag, and pitching moment coefficients and the key range and
1
C 2
C 2
L
endurance parameters, L C , D and L C , were analyzed over a range of Mach numbers and angles of
D
D
attack. The baseline model was tested with the initial configuration specified by AFRL/VA during the Fall 2006
semester. The follow-on effort conducted in the Spring of 2007 evaluated a first iteration of design optimization
which increased gap between the wings and varied decalage angles. The Joined Wing designs were preliminary –
no engine, landing gear or other external attachments were incorporated into the test models.
II. Background
The Joined Wing concept was developed by Ronald G.
Houck II, owner of Iron Hawk Enterprises LLC (IHE)2.
Mr. Houck received a patent and established a sponsor
through U.S. Representative David Hobson from the Ohio
7th District. The Air Force Research Laboratory Air
Vehicles Directorate agreed to perform evaluations on this
design because it provided a potential design concept for a
lightweight, long range, and high endurance UAV. Figure
1 is a schematic of Mr. Houck’s original design. The
design has since been modified for applicability in UAV
missions.
The potential aerodynamic benefit of this Joined Wing
design is a reduction in the amount of induced drag. To
completely eliminate the formation of induced drag, the
wing must have no tip. Figure 2 illustrates how the Joined
Wing design accomplishes this by joining the outboard
surfaces of the upper and lower wings.
A study was conducted at the University of Missouri
which compared biplane characteristics to mono-plane
characteristics4. This study compared the aerodynamic
performance of a monoplane and various biplane
configurations using the same wing area and average AR
while varying the span and chord. The University of
Figure 1: Joined Wing Design2
Figure 2: WPAFB Joined Wing Design
2
Missouri study determined the optimum biplane configuration would have a gap (vertical displacement of the upper
wing relative to the lower wing) of one chord length, a stagger (horizontal distance between the upper and lower
wing) of 0.875 chord lengths, and a decalage angle (difference in incidence angles between the upper and lower
wing, where positive angles result with a larger incidence angle on the upper wing) of -5°. The study concluded that
L
optimal biplane designs may produce a lower CD , increased D (indicating maximum endurance for jet propulsion
3
2
systems and maximum range for reciprocating engines), and increased L D (indicating maximum endurance for
reciprocating engines)4.
The Joined Wing combines the idea of the biplane design with the elimination of wingtips to minimize or
eliminate induced drag.
III. Set-Up and Procedure
A. Test Models
The original Joined Wing model was constructed by American Precision Prototyping, Tulsa, Oklahoma, using
stereo lithography. Its reference dimensions can be seen in Table 1. The model had approximately 1.98% wind
tunnel blockage at 0° angle of attack and 3.00% blockage at the maximum tested angle of attack of 20°. Figure 3
shows the baseline model. The moment reference center at which the pitching moment was resolved was 8.8125
inches from the back of the model (per AFRL/VAAA request). The zero angle of attack of the vehicle was defined
as the fuselage’s upper surface, located 0.75 inches from the nose, having a 3.62° incline from the horizon.
For comparison purposes, a model of the NASA Orbital Space Plane (see Figure 4) was used as an example of a
conventional mono-wing design. The Orbital Space Plane model was chosen because it had similar characteristics
to the Joined Wing model, as seen in Table 1. The wing span and total wing area (not planform areas) were similar
for both models.
Figure 3: Joined Wing Design
Model
Figure 4: NASA Orbital Space Plane
Model
Figure 5: Optimized Joined Wing Design
Model
The modified Joined Wing model was fabricated at USAFA and consisted of interchangeable parts that allowed
testing of two gaps (2.5 and 4.75 inches) and three decalage angles (0°, -1.5° and -3°). See Figure 5. All other
dimensions remained the same as the baseline model (refer to Table 1).
Wingspan (in)
Reference Area
(sq in)
Wingsweep, forward
wing (deg)
Wingsweep,
rear wing (deg)
Moment Reference
Center (in)
from back of model
Gap (in)
Decalage Angle (deg)
Table 1: Model Reference Dimensions
Joined Wing
Space Plane Optimized Joined Wing
18
17.6
18
160.0
142.9
160.0
21.8
49
21.8
-19.4
N/A
-19.4
8.8125
2.5
0
N/A
N/A
N/A
8.8125
4.75
0, -1.5, -3
3
B. Wind Tunnel
The investigations were conducted in the USAFA
Subsonic Wind Tunnel in the Aeronautics
Laboratory. The subsonic wind tunnel is a single
return, closed-loop type system with a 3ft x 3ft test
section. The maximum Mach number of the tunnel is
0.6. A schematic of the wind tunnel is shown in
Figure 6.
C. Test Section Configuration
The models were supported by an internal force
balance and an aft mounted sting / variable arc sector
in the test section. The arc sector was capable of
varying the angle of attack between ± 30°. This
assembly included a turn table able to rotate ± 26°
from the direction of the freestream flow to generate
sideslip angles. The wind tunnel test section and the
Houck Half Wing Aircraft mounted in the wind
tunnel are shown in Figure 7.
Figure 6: Subsonic Wind Tunnel5
Figure 7: Test Section Set-up6 (left) and Houck Half Wing Aircraft Mounted in Wind Tunnel (right)
An Able 100lb internal force balance was used to determine the forces and moments acting on the models. The
force balance measured normal, axial, and side forces along with a rolling moment.
D. Test Matrix
To determine the performance and static stability characteristics of the Joined Wing design, both longitudinal
and lateral-directional studies were conducted. For the longitudinal study, the aerodynamic coefficients CL, CD, and
Cm were obtained. The lateral-directional study evaluated the yaw and roll moment coefficients Cn and Cl with
variations in sideslip ( ). Since the longitudinal study provided the primary insight into aerodynamic performance,
only a longitudinal study was conducted on the Orbital Space Plane to provide comparisons with the Joined Wing.
Table 2 shows the test matrix for this experiment.
Initially, testing was planned for Mach numbers of 0.3, 0.4, and 0.5; however, due to large forces generated by
the model at high angles of attack and higher Mach numbers, the testing was limited to 0.3 Mach to keep the balance
loadings within the acceptable range for the 100lb force balance.
4
Table 2: Test Matrix, Baseline Model
Longitudinal
Mach
0.2, 0.25, 0.3
Alpha
-4° to 20°
(2° increments)
Beta
0°
Lateral-Directional
0.2, 0.25, 0.3
-4° to 20°
(2° increments)
-5° to 5°
(2.5° increments)
Space Plane
0.25, 0.3
-4° to 20°
(2° increments)
0°
For the optimization testing, focus was on the aerodynamic performance differences among the configurations,
best shown by a longitudinal study; consequently, only alpha sweeps at M=0.25 were conducted. Table 3 shows the
test matrix used for testing. As a final test done for comparison purposes, the upper wing was removed completely
to isolate the lower wing (this configuration was referred to as the half wing) and to discover its contribution to the
Joined Wing’s performance (see Figure 7). Note the truncated flow guide vanes resemble winglets. The same
planform area (160.0 sq in) was used for data reduction for comparison purposes.
Table 3: Test Matrix, Optimized Model
Baseline Design &
Increased Gap
Half Wing
Mach
0.25
Alpha
-4° to 18°
(2° increments)
Decalage Angle
0°, -1.5°, -3°
0.25
-4° to 18°
(2° increments)
N/A
IV. Results and Discussion
A. Uncertainty Analysis
As with all laboratory measurements, there exists some level of uncertainty in the results. This uncertainty arises
from inaccuracy inherent in the measuring devices and less than perfect repeatability of measurements. These
sources of error can be placed in two main categories—precision error (Px) which includes error associated with
repeatability, and bias error (Bx) which results from repeatable errors in measuring equipment. Precision error is
present when data is scattered around the true value, while bias error is present when all data points seem to be
shifted to one side of the true value. Bias error is a repeatable error and can be corrected using an offset value. The
overall uncertainty in the results was calculated from the contributions of both precision and bias error as shown in
Equation 15.
(
U = Bx2 + Px2
)
1
2
(1)
The important coefficients to consider for a complete analysis of the Joined Wing design were CL, CD, Cm, Cn,
and Cl. The coefficients were developed from the measurements taken from the force balance. Equation 2 presents
a typical formula for a lift coefficient.
CL =
L
=
qS
(N1 + N 2 )cos
Ax sin
(2)
1
P0
1
P0
P
1 (P0
P )S
Influence coefficients were used to account for the contribution of each measurement to the uncertainty in the
calculations. Influence coefficients are partial derivatives of a value with respect to a single measurement5. For
example, to isolate the contribution of one of the force balance normal forces (N1) to the calculation of the lift
coefficient, the partial derivative of CL was obtained with respect to N1.
5
Lift Coefficient, CL (-)
The uncertainty was calculated by
determining the ranges associated with both
2.5
bias and precision error, then multiplying them
by their influence coefficients. These bias and
2.0
precision values were then combined through
Equation 1 to calculate the total uncertainty.
The manufacturer’s published uncertainty for
1.5
the 100lb force balance was used to calculate
the bias error, and the precision error was
1.0
calculated from the standard deviations for
each measurement. Figure 8 shows the worst
0.5
case uncertainty calculations for the lift
coefficient versus angle of attack at a Mach
number of 0.25.
0.0
0
5
10
15
20
25
Although uncertainty was present for each
Angle of Attack, (deg)
of the aerodynamic coefficients, the range of
uncertainty was considered small, meaning
Figure 8: CL Uncertainty, M = 0.25
that the data obtained was reasonable and the
conclusions should be valid.
The small
uncertainty ranges were consistent for all flight conditions. Table 4 summarizes the absolute total worst case
uncertainty for each of the coefficient calculations.
Mach
CL
CD
CM
CN
CL
Table 4: Worst-case Uncertainty Analysis (Absolute Total Uncertainty)
Joined Wing Baseline
Space Plane
Lower Wing
Modified Gap (M=0.25)
0.2
0.25
0.3
0.25
0.3
0.2
0.25
0
-1.5
-3
Decalage
Angle
0.1387
0.2208
0.0810
0.0989
0.1024 0.0592 0.0556
CL
0.0528 0.0423 0.0377
0.0387
0.0616
0.0158
0.0320
0.0255 0.0129 0.0196
CD
0.0153 0.0108 0.0137
0.0368
0.0578
0.0520
0.1515
0.1568
0.0437
0.0310
0.0262
0.0411
0.0283
0.0072
0.0117
0.0025
0.0053
0.0055
The actual uncertainty was believed to be significantly lower than these values since the coefficients were
obtained by using the time average of one hundred data points taken during one second of testing at each position.
The time averaged data produced more accurate results for the coefficients since the major contributor to the
precision error uncertainty was the vibration of the model. Although the standard deviation of the one hundred data
points may have been significant due to model vibration, the vibration occurred about the true value, and averaging
the sample points should provide a good estimate of the true value with less uncertainty than estimated in Table 4.
B. Hysteresis
The effects of hysteresis were also examined. Hysteresis was noticed when angle of attack, as well as sideslip
angle, sweeps were performed. However, the amount of hysteresis in the data was considered small. Figures 9 and
10 present typical hysteresis data for the Joined Wing at a Mach of 0.25 for both the longitudinal and lateraldirectional studies. The hysteresis trends were consistent for all test cases.
6
Figure 9: CL vs
(Joined Wing, M = 0.2)
Figure 10: Cl vs
(Joined Wing, M = 0.25,
= 10°)
C. Lift Characteristics
An examination of lift coefficient data showed that the
Joined Wing design produced lift coefficients approximately
60% higher than those produced by the Orbital Space Plane
at all angles of attack, as shown in Figure 11. The lift curve
slope of the Joined Wing design was approximately
0.05 /deg which matches the lift curve slope produced by
the Space Plane. Mach number was found to have no
significant effect on the lift curve.
Figure 11: CL v
Comparison between Joined Wing
and Space Plane
0.35
0.30
Drag Coefficient, CD (-)
D. Drag Characteristics
A comparison between the drag polar of the Joined
Wing and the Space Plane is shown in Figure 12. The
Joined Wing had a minimum drag coefficient of 0.0925
which was 7 times higher than that of the Space Plane which
had a minimum drag coefficient of 0.0131. However, the
shallow nature of the drag curve seen for the Joined Wing
indicates that there was a possible reduction of induced
drag. Additionally, at higher values for the lift coefficient
(values above 0.55) the Joined Wing produced a lower drag
coefficient than the Space Plane, which indicates some
benefits in this lift regime. Mach number was shown to have
no significant effect on the drag coefficient.
0.25
0.20
Joined Wing
0.15
E. Longitudinal Stability Characteristics
0.10
The last coefficient analyzed in the longitudinal study
Space Plane
JW M = 0.25
was the change in pitching moment coefficient with angle of
SP M = 0.25
0.05
JW M = 0.3
attack. For a vehicle to be statically stable longitudinally, it
SP M = 0.3
0.00
must have a negative value for the slope, Cm . Figure 13
-0.4
-0.2
0.0
0.2
0.4
0.6
0.8
1.0
1.2
1.4
presents the pitching moment curve for the Joined Wing at
Lift Coefficient, CL (-)
varying Mach numbers.
Figure 12: Drag Polar Comparison between Joined Wing
At the low angles of attack (< 9°), the Joined Wing was
and Space Plane
unstable or neutrally stable; however, static stability was
obtained at the higher angles of attack. Additionally, the
graph indicated that the pitching moment never reached a value of zero meaning the Joined Wing was not capable of
trimming for the configuration and conditions evaluated. If the Joined Wing was flown in the test configuration, it
would experience a nose down pitching moment. Comparisons were not made with the Space Plane pitching
moment characteristics since the choice of moment reference center location would drive the comparison.
7
F. Lateral Stability Characteristics
The lateral stability of the Joined Wing was evaluated through
an examination of the rolling moment coefficient variation with
the sideslip angle. A stable vehicle in roll will have a negative
slope, Cl , meaning that when the aircraft experiences a sideslip,
it will generate a rolling moment to roll away from the sideslip.
Figure 14 presents rolling moment coefficient versus sideslip
angle for the Joined Wing at Mach 0.2 for varying angles of
attack.
According to Figure 14, the Joined Wing did not have
positive roll stability at low angles of attack. At a zero degree
angle of attack, the slope of the rolling moment coefficient curve
was positive, denoting a negative stability. As angle of attack
increased, the roll stability of the Joined Wing improved to its
maximum negative Cl value of -0.002/deg at 15° angle of
Figure 13: Cm vs
Curve for the Joined Wing
attack. These roll stability trends were consistent with little
variation between Mach numbers, each achieving the best
stability at a 15° angle of attack.
G. Directional Stability Characteristics
An examination of the yawing moment coefficient with
respect to the sideslip angle was conducted to evaluate directional
stability. A vehicle stable in yaw experiences a positive slope,
C n , when plotting yawing moment coefficient and sideslip
angle.
A positive
Figure 14: Cl vs
for the Joined Wing (M = 0.2)
Figure 15: Cn vs
for the Joined Wing (M = 0.2)
C n means that the when the vehicle
experiences sideslip, the vehicle will develop a yawing moment
which will redirect the nose back into the wind and align the
body with the freestream. Figure 15 presents the yawing
moment coefficient characteristics for the Joined Wing at a
Mach number of 0.2.
It can be seen that the Joined Wing had a positive C n at
zero angle of attack. As angle of attack increased, the joined
wing became nearly neutrally stable at 5° and 10° angle of
attack, and then unstable at 15° and 20° angle of attack. At an
angle of attack of 15°, the Joined Wing had the worst yaw
stability with the largest negative slope. Comparisons between
Mach numbers yielded consistent behavior, except at the zero
angle of attack with no sideslip, for which the model was stable
for Mach 0.2 and 0.3, but near neutrally stable for Mach 0.25.
H. Endurance and Range Characteristics
Endurance and range potentials are
dependent on the propulsion system. The maximum range potential for a jet
1
2
engine is indicated by the ratio C L C , the maximum endurance potential for a jet propulsion system and the
D
maximum range potential for a reciprocating engine is denoted by the
3
L
D ratio, while the maximum endurance
potential for a reciprocating engine is found through the ratio C L C . Figure 16 shows that the range potential
D
2
1
with a jet engine is far inferior to that of the Space Plane. The Space Plane produced a maximum C L C that was
D
2
8
3.15 times higher than that of the Joined Wing. However, at the high lift coefficients (above 0.55) the Joined Wing
had better performance.
25
20
JW M = 0.25
SP M = 0.25
JW M = 0.30
SP M = 0.30
Space Plane
CL
1/2
/CD
15
10
Joined Wing
5
0
-0.4
-0.2
0.0
0.2
0.4
0.6
0.8
1.0
1.4
1.2
Lift Coefficient, CL (-)
Figure 16: Max Range for Jet Engine Comparison
between Joined Wing and Space Plane
6
The potential for endurance with a jet engine and range
with a reciprocating engine is shown in Figure 17. The
L
Joined Wing proved to have an inferior D ratio compared
Joined Wing
5
Space Plane
3
CL
3/2
/CD
4
2
JW M = 0.25
SP M = 0.25
1
JW M = 0.30
SP M = 0.30
0
-0.4
-0.2
0
0.2
0.4
0.6
0.8
1
Figure 17: Max Endurance for Jet/Range for
Reciprocating Engine Comparison between
Joined Wing and Space Plane
1.2
Lift Coefficient, CL (-)
Figure 18: Max Endurance for Reciprocating
Engine Comparison between Joined Wing and
Space Plane
to the Space Plane by approximately 78%. However, the
potential for the Joined Wing was illustrated by better
performance at extremely high (>13°) and low (<2°) angles
of attack.
Lastly, to evaluate
endurance potential for a reciprocating
3
2
engine, the ratio C L C was examined. Figure 18 shows the
D
comparison for the Joined Wing and 3the Space Plane.
2
The Space Plane produced a C L C ratio, indicative of
D
endurance potential for a reciprocating engine, that was
approximately 8.3% higher than the maximum produced by
the Joined Wing. However, the Joined Wing out performed
the Space Plane for lift coefficients above 0.55.
I. Optimized Model
The optimized Joined Wing evaluation varied gap
(baseline=2.5 in; modified=4.75 in) and decalage angle (0°, -1.5°,
-3°). Figure 19 shows the longitudinal static stability derivative.
Because all slopes are similar and negative, all configurations had
similar positive longitudinal stability.
Figure 20 shows the results of the coefficient of lift tests at
M=0.25 with varying gap and Mach number. The increased gap
and lower-wing only (half wing) configurations produced similar
lift coefficients. The baseline gap produced significantly lower
lift coefficients, probably due to interference of the lower wing to
the flow around the upper wing because of the small gap. With
increasing decalage angle, the difference between lift coefficients
produced by the modified and baseline gap configurations
decreased.
Figure 21 shows the drag polar comparison for all
combinations of gap and decalage angle as well as the lower-
Figure 19. Longitudinal Static Stability
9
wing only configuration. The lower-wing only configuration produced the best results although it is interesting to
observe the decrease in induced drag with increased gap.
0.35
Drag Coefficient, CD (-)
0.30
0.25
0.20
Base Dec 0
Base Dec 1.5
Base Dec 3
Mod Dec 0
Mod Dec 1.5
Mod Dec 3
Half
BASE
0.15
MOD
0.10
HALF
0.05
0.00
-0.4
Figure 20. Lift Coefficient Comparison
(Decalage Angle=0°)
-0.2
0.0
0.2
0.4
0.6
Lift Coefficient, CL (-)
0.8
1.0
Figure 21. Drag Polar Comparison
The key performance parameters analyzed were those that show maximum range and endurance for jet and
reciprocating propulsion systems. Figures 22, 23, and 24 show these results. These plots show that increased gap
was a significant improvement over the baseline design. However, the lower-wing only configuration consistently
had the best performance.
25.0
6.0
HALF
10.0
4.0
3.0
MOD
Base Dec 0
Base Dec 1.5
Base Dec 3
Mod Dec 0
Mod Dec 1.5
Mod Dec 3
Half
HALF
3/2
MOD
BASE
CL
CL
1/2
/CD
15.0
5.0
/CD
20.0
Base Dec 0
Base Dec 1.5
Base Dec 3
Mod Dec 0
Mod Dec 1.5
Mod Dec 3
Half
2.0
5.0
1.0
BASE
0.0
0.0
-5.0
-0.4
-1.0
-0.2
0.0
0.2
0.4
0.6
Lift Coefficient, CL (-)
0.8
1.0
1.2
Figure 22. Max Range for Jet Engine Comparison
-0.4
-0.2
0.0
0.2
0.4
0.6
Lift Coefficient, CL (-)
0.8
1.0
1.2
Figure 23. Max Endurance for Reciprocating
Engine Comparison
Generally, decalage angle had little effect on the test
results. When analyzed closely, a decalage angle of -1.5°
was consistently the best of the three angles, suggesting it
is the optimum decalage angle.
V. Conclusions
Figure 24. Max Endurance for Jet Engine / Range
for Reciprocating Engine Comparison
Overall, the baseline Houck Joined Wing was not
found to provide any significant advantages as a UAV
design concept. The Joined Wing produced high values
of C D , with a minimum drag coefficient 7 times higher
than that produced by the Space Plane. However, the
drag polar had a fairly shallow slope, indicating the
possibility that the Joined Wing met its objective of
decreasing induced drag. At lift coefficient values above
0.55, the Joined Wing produced lower drag coefficient
10
values than the Space Plane.
The Joined Wing’s longitudinal stability tended to improve with increasing angle of attack. The Joined Wing
was longitudinally unstable at angles of attack below 9° and stable at angles of attack above 9°. However, the
pitching moment curve never crossed the y-axis, showing that the Joined Wing was unable to trim for the test
conditions evaluated. The Joined Wing’s roll stability increased with angle of attack, with it being the most stable at
15° angle of attack. Neutral yaw stability was achieved at 5° and 10° angle of attack and positive yaw stability
existed at 0° angle of attack. However, the Joined Wing was unstable at 15° and 20° angle of attack.
The primary focus of this effort was to 1determine the Joined Wing’s capabilities as a long range, long endurance
2
UAV. The Space Plane had a higher C L C ratio for maximum range for a jet engine until the lift coefficient
D
reached 0.55, at which point the Joined Wing out performed the Space Plane. At high (>13°) and low (<2°) angles
L
of attack, the Joined Wing had a better D ratio, meaning it had improved maximum endurance potential for a jet
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engine and range for a reciprocating engine. The parameter that showed the most potential was the C L C ratio,
D
2
which indicates the maximum endurance for a reciprocating engine. The Joined Wing had higher values than the
Space Plane at lift coefficients over 0.55.
The optimization effort proved successful, with the variation in gap having much more significant effects than
varying the decalage angle. Increased gap significantly improved aerodynamic performance. However, the lowerwing only configuration performed better in all areas than any configuration of the Joined Wing design, suggesting
that the upper wing is inhibiting rather than helping the aircraft’s performance.
VI.
Recommendations
Additional testing should be accomplished in the form of water tunnel testing, CFD analysis, pressure sensitive
paint and/or oil tests, to determine if the upper wing is contributing to the Joined Wing’s performance. If it is not,
the model’s reconfiguration should be focused on making the upper wing a contributor to the design—probably by
using a negative decalage angle and a conventional taper ratio.
The Joined Wing produced some interesting results. Its optimization in the areas of gap and decalage angle
showed improvements in aerodynamic performance that may warrant further optimization. The unique design and
its possibilities should be further studied to find out how useful this concept can be and how to improve upon its
strengths.
Acknowledgments
Wright-Patterson AFB/VAAA for funding
Mr. Ken Ostasiewski for wind tunnel assistance
Mr. Brian Kinsey for model definition
Dr. Post for paper revisions
Capt Elaine Byant for program guidance
Dr. McLaughlin for uncertainty analysis spreadsheet
References
1. Brandt, Stiles, Bertin, and Whitford, Introduction to Aerodynamics: A Design Perspective, United States
Air Force Academy, AIAA.
2. Fry, Timothy Ph.D., Phase One Report – Houck Configuration, University of Dayton Research Institute,
Dayton, OH, Aug 2006.
3. Lennon, Andy. Model Airplane News. “Biplane design,” Air Age Publishing, February 1998.
4. Olson, E. Carl and Selberg, B. P., Experimental Determination of Improved Aerodynamic Characteristics
Utilizing Biplane Wing Configurations, University of Missouri-Rolla, Rolla, MO, Journal of Aircraft, Vol 13, No 4,
1976.
5. Ken Ostasiewski. United States Air Force Academy Department of Aeronautics.
6. Batish, Paul, C1C, and C1C Johnson, Kip. 471 Report: X-38. USAFA. Fall of 1999.
7. Wheeler, Anthony and Ganji, Ahmed, Introduction to Engineering Experimentation, 2nd ed. Pearson
Education, Inc, Upper Saddle River, NJ, 2004, pp. 180-210.
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