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ROLLED-UP VORTEX DYNAMIC NUMERICAL STUDIES OF BIRD'S-BODY-TYPE FIGHTER

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International Journal of Mechanical Engineering and Technology (IJMET)
Volume 10, Issue 03, March 2019, pp. 304-316. Article ID: IJMET_10_03_032
Available online at http://www.iaeme.com/ijmet/issues.asp?JType=IJMET&VType=10&IType=3
ISSN Print: 0976-6340 and ISSN Online: 0976-6359
© IAEME Publication
Scopus Indexed
ROLLED-UP VORTEX DYNAMIC NUMERICAL
STUDIES OF BIRD'S-BODY-TYPE FIGHTER
Sutrisno
Department of Mechanical and Industrial Engineering, Faculty of Engineering, Universitas
Gadjah Mada, Yogyakarta, Indonesia 55281
Setyawan Bekti Wibowo
Department of Mechanical and Industrial Engineering, Faculty of Engineering, Universitas
Gadjah Mada, Yogyakarta, Indonesia 55281.
Department of Mechanical Engineering, Vocational College, Universitas Gadjah Mada,
Yogyakarta, Indonesia 55281.
Sigit Iswahyudi
Department of Mechanical and Industrial Engineering, Faculty of Engineering, Universitas
Gadjah Mada, Yogyakarta, Indonesia 55281
Department of Mechanical Engineering, Faculty of Engineering, Universitas Tidar, Magelang,
Indonesia 56116.
Tri Agung Rohmat
Department of Mechanical and Industrial Engineering, Faculty of Engineering, Universitas
Gadjah Mada, Yogyakarta, Indonesia 55281,
ABSTRACT
The fuselage type of a fighter greatly influences the aerodynamic characteristic
especially in vortex generating which determines the agility and maneuverability. This
paper elucidates the study of the vortex dynamics of a fighter with a bird's body type
fuselage (BBTF), at constant canard deflection angle, using computational fluid
dynamics (CFD). The CFD study employs Q-criterion to probe vortices and structured
logarithmic grid to maintain the micro-gridding effectiveness of the turbulent
boundary layer. The flow visualizations are coupled with shear-wall streamline to
produce complete vortex dynamic pattern. The results show that the configurations of
the bulky-head aerodynamic, the body, canard and main wing of the BBTF affect
vortex core trajectories on the blended-body-main wing and the canard that very close
to the surface, which gives higher coefficients of lift, of order 1.40 at a high angle of
attacks causing its high agility and maneuverability. The numerical calculations show
that there is no negative surface pressure distribution (SPD) on vertical walls near the
canard and wings of BBTF fighters. On other design, negative SPD on vertical walls
waste a lot of fuel energy, including weakening the lift coefficient.
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Keywords: Fighter aircraft, CFD, vortex dynamic mechanism; fuselage effect; Qcriterion.
Cite this Article Sutrisno, Setyawan Bekti Wibowo, Sigit Iswahyudi and Tri Agung
Rohmat, Rolled-Up Vortex Dynamic Numerical Studies of Bird's-Body-Type Fighter,
International Journal of Mechanical Engineering and Technology, 10(3), 2019, pp. .
304-316.
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1. INTRODUCTION
The difference between a passenger plane and a fighter is the assumption of its flight speed
operation. Passenger aircraft are characterized by subsonic to transonic flying speeds, while
fighter aircraft preferably work reliably at supersonic speeds. So, the configuration of those
planes is different. The passenger plane has a low swept wing, so that the airflow around the
plane is of a straight flow with little mixed with the rolling vortices, whereas on fighter
aircraft with medium to the high swept wing, so the airflow around the fighter is more mixed
and mostly dominated by rolled-up vortices.
On passenger aircraft, the wind velocity energy is converted into the high-velocity flow to
form a low-pressure zone above the wing airfoil surface to create lift force on the wing, tail,
and control surfaces as the role of vortex swirls, in passenger aircraft, is lesser. Compared to
combat aircraft, the force developments are more dominated by the role of vortices, yet
straight flow is still significant. The role of the vortices start above the wing brought up by the
rolled-up-vortex effect, which also then plays a lot on the canard or leading edge extension
(LEX) of the fighters.
At present, the widely known fighter flow energy transmission is the rolled-up vortex
(RuV) effect, such as above the wings, LEX, and canard. The fuselage effects are still not
well known, mainly due to the effects of the blended wing-body, the blended canard-body,
including the effect of the bulge of the head and the neck, canard, wing's height difference,
planar distance, and the flat horizontal wing-surface from head to tail. The fuselage effect is
due to the head protrusion as the head has an airfoil-like cross-section that its 3-D effect
produces an airfoil lift as the air flow above the head is much faster than that at the base.
What is the difference between the fuselage and a rolled-up vortex effect?
The rolled-up vortex is the trailing vortex wake behind the lifting wing or delta wing,
which is the natural tendency of free vortical shear layers to roll up into well-defined vortices
with the essentially inviscid outer flow rotating around an inner viscous core. CFD or
mathematical analysis is the means for calculating the characteristics of the complete flow
field over the wing or delta wing as well as its wake and the rolled-up tip vortices. It is
renowned that the vortex layer, formed by the union of upper and lower wing surface
boundary layers, behind the trailing edge of a running and delta wing, rolls up into two vortex
cores. What is the fuselage effect?
The fuselage effects are the effects generated by any form of body or obstruction,
including laminar stream form, which settles in the path of the flow so that the curved contour
of the obstruction cleaves the flow apart. If the obstruction is aerodynamic enough, it might
generate special RuV effects. Scientists described, they found the ocean wind passes through
the island. The curved contour of the island cleaves the ocean wind flow apart; the RuV
effects generates strong vortex cores following the site of the island [1]. Similar things happen
to the bifurcation of the vortex cores on the backward wind turbine blades, for blunt-blades
generally have potent vortex-tip, at the same time also generate a second weak vortex center.
This second vortex center appears due to the rolled-up vortex effect as a result of the swept
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effect in the mid-zone of the backward wind turbine blades. Laminar zone deflects the flow
and creates a rolled-up vortex effect.
Several scientists explored the rolled-up vortex effect on different wings. Sohn et al. [2]
investigated the effect of a center body on the vortex flow of a double-delta wing with leading
edge extension. Heyes and Smith [3], learned in depth the modification of a wing tip vortex
by vortex generators. Niu et al., [4] examined numerical research on the effect of variable
droop leading-edge on oscillating NACA 0012 airfoil dynamic stall, and Abene and Dubois
[5] observed fundamental aspects of the vortex flow on cones. Researchers have conducted
several highly advanced types of research in fighter vortex dynamics. The growth of
aerodynamic efficiency and the growth of the wing structural stress of a hyperbolic winglet
concept on DLR-F4 Aerodynamics are studied for typical transport aircraft wing-body, after
installing classical Whitcomb winglets of different configurations and a delta wingtip fence,
revealing the interaction of sub- and transonic airflow dynamics. A multidisciplinary winglet
efficiency estimation criterion is suggested in real flight conditions [6].
Further study for future aircraft design, including military purposes, could develop chief
advancement of aircraft technology. Flight physical aspects and methods of future military
aircraft designs have been thoroughly explored [11]. Schminder has carried out a feasibility
study of different methods for the use in aircraft conceptual design [12]. Vortical and
turbulent structures for an aircraft at drift angle 0, 12 and 30 degrees have been learned [13].
A study of an asymmetric vortices synergy over chinned forebody/wing configuration at high
angles of attack (AoA) has shown that forebody vortices can postpone breakdown of wing
vortices [14]. Zhang et al. have investigated numerically the characteristics between the
canard and wing of the canard-forward swept wing aircraft configurations at different canard
positions focusing on the interference between the canard and wing [15].
Some researchers have performed numerical model simulations of F-16XL fighter.
Boelens et al. have executed geometry and computational grids with structured grids and the
unstructured grids used in cranked- arrow wing aerodynamics project [16]. Lofthouse and
Cummings have conducted numerical simulation at flight-test conditions using delayed
detached-eddy simulation [17]. Some simulations at flight conditions using hybrid near-body
computational fluid dynamics have been carried out as well [18]. Nevertheless, the
application of the Q-criterion algorithm so far has not been combined with the footprint of the
streamline for the vortex development, to scrutinize the complete picture of the vortex
dynamics. According to the fuselage effect theory [19], fighters fuselage configurations give
different spreading effect on the fighter wing, determining the lift force generated.
Following the fuselage effect theory, canard fighters have two different fighter types,
namely a) the bird's body type fuselage (BBTF) fighter, as the head and the neck of the
fighters resemble the head and the neck of a bird and b) the straight body type fuselage
(SBTF) fighter, as most canard fighter bodies have straight fuselage. The BBTF fighters
include Sukhoi Su-30, Sukhoi Su-27, and F-35 Lightning. Several advanced scientists have
explored numerical investigation to some fighters. Boelens have conducted CFD analysis of
the flow around the X-31 aircraft at a high angle of attack [20]. Chen et al. have studied the
effect of sideslip on high-angle-of-attack vortex flow over a close-coupled canard
configuration [21]. Ghoreyshi et al. have performed simulation validation of static and forced
motion flow physics of a canard configured TransCruiser [22]. Ghoreyshi et al. have learned
transonic aerodynamic load modeling of X-31 aircraft pitching motions [23]. Schütte and
Rein have examined experimental and numerical aspects of simulating unsteady flows around
the X-31 configuration [24]. In this paper, the Q-criterion algorithm is employed, so that the
footprint of the streamline for the vortex development and the wall shear is combined to
scrutinize the complete picture of the vortex dynamics.
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For that purpose, we perform some tests, namely a) mesh independence test in the form of
y+ value selection, beyond that although with grid reduction, the error of computation will not
change because it has reached the accuracy of Kolmogorov microscale. Then, b) we consider
the decrease of gridding scale logarithmically according to the distribution of turbulence
speed, c) selection of turbulence equation models k-θ, k-ω, or SST, which take into account
the practicality of grid values and appropriate accuracy.
It is essential d) to display the global vortex dynamics with streamline and Q-criterion to
illustrate the strength of the flow and vortex in detail, and then e) it gives quantitative results
that describe the strength of the flow and vortex swirls.
The Q-criterion algorithm in CFD research is applied to analyze vortex especially vortex
cores along the vortex, the speed of the axial vortex Uc / U∞ (z) relative to the local free flow
velocity and Cp pressure coefficient along the center of the vortex, surface pressure
distribution (SPD) in the left and right fighter wing surface, as well as the vortex core
trajectory height. The magnitude and direction of the vortex core are measured using QCriterion with the equation:
[| |
| | ]
According to the flow identification of the Q-criterion, the vortices of the incompressible
flow are identified as fluid regions connected to the positive second invariant of the tensorgradient tensor ∇u, where ∇u = S + Ω, S is the tensor-rate tensor, and Ω is a vorticity tensor.
In this simulation, fighter speed 0.3M suits the speed of dogfighting battles that have to be
done at slow speeds. This is in accordance with the requirements for vortex identification
criteria which applies only to incompressible flows. In this investigation, the Q criterion was
chosen as the vortex identification criterion which only applies to the in-compressible flow,
that is the flow at Mach number 0.3. At higher speeds, M≥ 0.3 is avoided because flow creates
drag divergence .
The purpose of this investigation is to study the fuselage effects of aircraft on the vortex
dynamics of BBTF fighters. The results of BBTF fighter design analysis, using a structured
logarithmic mesh, apply the numerical simulation methodology in a section of results, to
maintain the accuracy of the turbulent boundary layer on a micro scale.
2. RESEARCH METHODS
2.1. Numerical and Vortex Dynamics Analysis Methods
The BBTF fighters included Sukhoi Su-30, Sukhoi Su-27, and F-35 Lightning. In this study,
the authors chose one of these fighters as a representative of BBTF fighters, as shown in
Figure 1., with some simplification in symmetrical models, and some detailed drawings, such
as an antenna. A net on the Sukhoi Su-30-like model aircraft was created by identifying the
aircraft parts and then dividing them into several blocks based on surface changes of the
aircraft. By changing the net size, we structured the hexahedral mesh nets, starting from the
wall portion as the smallest size and enlarging logarithmically to the outside [20].
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(a) Model of the BBTF fighter CFD model
(b) computational domain and mesh for the
model (symmetrical)
Figure 1 The dimension of the BBTF fighter CFD model and mesh configuration of computational
domain and boundary condition
In this study, one should perform a mesh independence test to obtain an accurate
computational model. The test results showed the convergent value occurred at the number of
cells starting at 5 million. This research uses cell number as much as 6,012,908 (~ 6 million).
The y+ = 1 was used to determine the smallest size of the cells near the wall, with the lowest
cell was 0.0069 mm. Since the dog-fighting is done at slow speed, the aircraft action is
completed at a speed of 0.3M. Therefore, the vortex dynamics pattern around the aircraft
would be symmetrical, so to save time, the computation is carried out with the half model.
Figure 1.b and 1.c show the shape of the net. The domain of the computation was half boxshaped the model that would be formed symmetrically. The boundary conditions were
determined as velocity inlet, pressure outlet, and the symmetrical planes.
This research involved variations of the AoAs from 200 to 600. The inlet velocity was set
at 0.3 M (114. m/s) streaming on the plane surface at an intensity of turbulence of 0.08%. The
flow model was on the Navier-Stokes equation employing the numerical simulation using the
finite volume method.
Vortex dynamics analysis is used to evaluate the fuselage and RuV effects of the BBTFlike fighter canard. Vortex dynamics analysis involves flow visualization to analyze fighters
and a review of the measurement results. Flow visualization consists of steps to plot the
limiting streamline, drawing footprints from time to time. The plot of flow visualization of the
vortex core is also presented, which may also generate the second vortex core. Afterward, the
measurement results are analyzed. The first result is related to the strength of the vortex
center, which is the axial vortex core velocity relative to the local freestream velocity
Uo/U∞(z) and the coefficient of pressure (Cp) of the vortex core. The trajectory height and
spanwise location of the vortex center are then measured. To demonstrate the scale of the
vortex core one uses the Q-Criterion.
2.2. Numerical Validation
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The wind tunnel experimental measurement data, used for validation, is a result of an X-31
experiment with high angle-of-attack aircraft configuration. The model data was used at first
by Boelens [20], from the DNW Low-Speed-Wind-Tunnel (LSWT) Braunschweig, DLR. The
test VN01004 has been chosen from the LSWT data set as a reference for the aerodynamic
force data. This test run constitutes angles of attack α ranging from −60 to 550. The wind
tunnel velocity has been 60 m/s (0.18 M).
(a) Lift coefficient (CL)
(b) Drag coefficient (CD)
Figure 2 CL and CD for X-31 CFD model employing the k-ω, k-ε and SST turbulence equations
versus an experiment, done at DNW, LSWTB, DLR [20] for varieties of AoAs
Figure 2 shows the comparison between this experimental data against the lift and drag
coefficient CL and CD for X-31 CFD model employing the k-ω, k-ε, and SST turbulence
equations. The calculation results of the Navier-Stoke equation, utilizing Fluent software,
conducted to CFD models of X-31, were versus Boelens' X-31 DNW experiment. It shows
that in comparison with the model k-ε, and SST. The model with SST gives the best match.
3. RESULT AND DISCUSSION
3.1. The Results of The CFD Calculations
b)
c)
Figure 3 Pictures show: (a) Lift coefficient, CL for BBTF CFD models for different AoAs (α)
= 00 to 600 for constant canard deflection (c = 00) compared with the water tunnel (WaTu)
experiment [30]. The better blended-wing-body design caused (b) insignificant negative SPD
a)
on the walls above the canard, the main wing leading edge and c) front view of BBTF fighter
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Figure 3 shows the lift coefficient for BBTF-like CFD model for a different angle of
attack: (α) = 00 to 600. The results are compared to previous studies using the Water Tunnel
[30]. For AoA = 00 to 400 the lift coefficient curves increase linearly. But, in Figure 3, at the
AoA = 440 - 450 and higher, the lift coefficient curves experience a spike surge due to the
bursts of vortex core from above the canard, bumping into the wing leading edge. Therefore
we limit our research up to AoA = 600. Further research needs to rearrange the canard AoA
(αC). With the canard c rearrangement in Figure 3, the lift coefficient, CL for BBTF-fighter
could be around 1.4. It means that the BBTF fighter would have high agility and
maneuverability.
3.2. Vortex Cores Visualization
(a) vortex core at AoA 300
(e) total pressure loss and wall shear at AoA 400
(b) vortex core at AoA 400
(f) total pressure loss and wall shear at AoA=500
(c) vortex core at AoA=500
(g) streamline at AoA 300
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(d) total pressure loss and wall shear at AoA
300
(h) streamline at AoA 400
Figure 4 Flow visualization of the canard, and the main wing vortex cores, total pressure loss, wall
shear flow, and the corresponding streamlines.
The qualitative trajectory of the vortex cores shown on AoA = 100 to 400 was very close
to the canard and wing surfaces. Figure 4 displayed the vortex structure in details for AoA, α
= 100 to 500, using the Q = 2.5 × 105 s−2. Following the fuselage effect theory [19], in Figure
4, one could observe the wing vortex core. They denote their vortex cores at a different AoA,
(α) = 100, 300, 400, and 500. Figure 4 shows the flow visualization of 1) the canard and wing
vortex cores and the flow configuration of the wall pressure. The colored circles denote the
vortex strength. White lines present the trajectory of the vortex cores through the circles. As
the fighter moves forward, the fuselage deflects the canard flow, so the curved head contour
cleaves the flow apart. It strengthens the RuV effects above the canard. The AoA increases to
(α) = 400, the vortex core strengthens, thereby generating canard lift. Figure 4, likewise,
displays 2) the trajectory of vortex core at varies the different AoA, (α) = 10 0 to 400, and
presents, as well, 3) the adverse surface pressure distribution (SPD) at the wing, showing
prime vortex re-attachment.
Figure 4 does not show significant negative surface pressure distribution (SPD) at the
vertical wall of the BBTF-like-fighter, close to the canard and the wing, as seen in Fig. 3.b..
Negative surface pressure distribution (SPD), on vertical walls near the canard and wings,
would waste a lot of fuel energy, including weakening the lift coefficient creation. This
energy saving is due to the implementation of the blended-wing body design and the
alignment, the same height between the canard and the wing leading edge. The qualitative
trajectory of the vortex cores shown on AoA = 100 is very close to the surfaces of the canard
and wing. According to fuselage effect theory, for (α) = 00 to 400, the trajectory vortex core in
the canard and on the main wing is very close to the surface, creating strong CL.
3.3. Flow Measurement at AoAs = 100 to 600
Figures 5 shows the thermodynamic process clearly. It begins with mechanical energy on the
wing and canard leading edge, turns into kinetic energy of vortex core shown in Figure 5.a,
and turns into potential energy of the negative pressure along the vortex core shown in Fig.
5.b. The induced surface-mechanical energy changes to a negative SPD, whose crossdistribution is shown in Figure 5.c, at 60% of the chord, and in Figure 5.d, on 30% of the
chord. Both figures describe the distribution of lift forces generated.
Here the figures illustrate the fuselage effect, showing a) the influence of the fighter head
and neck bulges, emitted as in Fig.5.a. , b) the influence of vortex core bursts onto the wing
leading edge, and c) the effects of the transverse wing planes, smooth elongated surface free
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of slight bumps, as in Fig.5.c. and 5.d. Moreover, the essential cause is due to blended-wingbody design, that there is no negative SPD on the vertical wall surface, as in Fig.3.b.
Figure 5.a describes axial velocity variations along the vortex cores, Uc/U∞ (α), for
various AoA from 100 to 600. For AoA = 100, 150, initially held at a low velocity, then rises to
a maximum of about 2, the back is weak even though for AoA = 150 is still around 1. For
AoA = 200, 300, and 400 the velocity directly increases to a maximum of 2.3. For AoA = 500
and 600, the flow velocity begins to fluctuate with axial velocity ranging around 1. It shows
that the vortex core flow is too close to the wing surface. Figure 5.b illustrates the pressure on
the vortex cores, which shows the efficient conversion of mechanical energy so that at AoA =
100 to 400, represents the low mechanical energy remaining. While at AoA = 500, 600 the
percentage of energy conversion becomes mechanical energy increases, the height of the
vortex core trajectory is too low.
(a) the ratio of velocity variation Uc/U∞(α)
(b) the coefficient of pressure variation Cp (α)
(c) distribution of surface pressure at 60% wing
root chord
(d) surface pressure distribution at 30% wing
root chord
(e) the height of wing vortex core (VCTH)
(f) ) spanwise location of the main wing of the
vortex core
Figure 5 Graph of the velocity ratio variation, pressure coefficient surface pressure distribution height
and span-wise location of wing vortex core for different AoAs (α) at AoAs (α) = 100 to 600.
Figure 5.c shows SPD above the main wing in 60% of the root chord. In the middle of the
left wing and the middle of the right wing, for AoA = 100, 150, 200, 300 and also 500, 600, the
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SPD potent, -Cp = 6, while at AoA = 400, the SPD = 8, is strong. At the middle of the spine,
the weak SPD at AoA = 100, 150, 200, 300 and also 500, 600, -Cp = 2-4 being on AoA = 400,
SPD = 6, is very strong. And on the left edge and the right edge of wing, for AoA = 10 0, 150
and 600, SPD = 1-2.5, is weak, at AoA = 200, 300, 400 the SPD potent, -Cp = 4-8, while at
AoA = 500, SPD =12, is extra strong. Figure 5.d shows SPD above the main wing in 30% of
the root chord.
In the entire cross-section, the SPD values for all AoA are almost the same, decreasing
slightly compared to Figure 5.d. Slight changes in AoA = 500 slightly jiggled SPD edged
around 10, and for AoA = 600 SPD on edge come jiggle at about 6. Figure 5.e and 5.f
illustrate that in AoA up to 500 - 600, it is necessary to adjust the canard deflection to
overcome the trajectory of the fluctuating vortex core height. By adjusting the canard angle
the - SPD would rise high for higher fighter AoA.
3.4. Discussion
For the case of BBTF fighter, a grid independence test is capable of capturing vortex
dynamics on the wall. They report Δx + 1 / Δx, which is logarithmically structured, so it takes
only 5 million mesh, which is the natural profile of the turbulent boundary layer. Results of
CL and CD numerical values for α from 0 to 600 are displayed. With SST turbulence model
obtained error = 0.5 - 1.0%. This BBTF fighter reaches very high CL, due to 1) BBTF fighter,
due to fuselage effect as the head and the neck resemblance the head and the neck of a bird, 2)
it has flat wing from-head-to-tail, and 3) it has a similar height between canard and its wing.
From the above description, one can conclude that the negative impact of RuV and the
fuselage effect on the BBTF fighters are in the form of reduced C L and CD increase. RuV and
fuselage effects strengthen the CL increase depending on the configuration. This fuselage
effect on BBTF fighter occurs in the following, due to the effect of a) blended-wing-body
design, b) the same height between the canard and the wing leading edge, c) the fighter bulges
on head and neck of BBTF fighters, d) the vortex core bursts onto the wing leading edge, e)
the transverse wing elongated surface planes. Trajectory vortex cores above the canard and
the main wing are very close to the surface; the vortex cores could give high C L, and be easily
tangled to make the velocity fluctuating at high-SPD, which needs canard angle adjustment.
On the BBTF fighter, there is no negative SPD found on the vertical wall surface.
Moreover, by adjusting the canard angle on BBTF the - SPD would probably still close to the
wing surfaces even for higher fighter AoA. Therefore, up to AoA = 400, overall CL is very
high. As vortex core split on the main wing much energy absorbed into -SPD gives high CL,
and decreased the end power of vortex cores. At the angle of AoA of 100 and 500 to 600, the
vortex cores fluctuate so that resetting canard deflection is required. The arrangement above
needs careful research for BBTF fighters. Figure 2 presented several examples of the
combined effect of CL and CD for X-31 fighter, and in Figure 3.a delivered C L for BBTF
fighters.
In the following are shown some of the results of previous studies compared to the results
of the BBTF fighter's research. Compared to the Sukhoi Su-47 [25], the Su-47 has a higher
lift-to-drag ratio, greater air battle maneuvering capacity, higher range at subsonic speeds. In
accordance with the Sukhoi BBTF research in water tunnel [26], the study emphasized the
increase in the maximum coefficient of lift due to the effect of the aircraft body. Compared
with the results of the Saab JAS Gripen C [27] study on AoA around 400 maximum CL values
of Gripen reached 1,431 due to the effect of the slotted wing. The maximum C L value of
Eurofighter [28], Chengdu J-10 from CFD computation [19] and water tunnel measurement
[29] is lower than the BBTF fighter.
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4. CONCLUSION
This study reports the RuV and fuselage effects of BBTF fighter. The BBTF fighter has dense
configurations in lift, maneuverability, and agility.
In visualization with BBTF fighter, one could find as follows: There was no -SPD on the
vertical wall, most probably due to intensive utilization of blended-wing-body design.
Trajectory vortex cores above the canard and the main wing are very close to the surface, so
the vortex cores could a) give high CL up to 1.4, or b) be easily tangled to make the velocity
fluctuating at high-SPD. Measurement of BBTF fighter vortex dynamics products indicates
that the transformation into -SPD gives high CL so that at the end the power of vortex cores
strongly decreases. At the angle of AoA of 10º and 50º to 60º, the vortex cores fluctuate.
Moreover, the vortex core is close to the wing surface, above the back-flat-wing surface
from head to tail, makes -SPD over the rear wing stronger and more evenly distributed. The SPD on the back wing is quite high, especially on the edge of the wing, the highest is at the
30% rear wing.
ACKNOWLEDGMENTS
The authors would like to express heartfelt gratitude to Dr. Bramantyo for a fruitful session,
useful suggestions, and collaboration. We appreciate the help of our students Wega, David,
Patricius, and Yogi, and the lab staff members, Ponimin and Wajiono, for giving their help in
construction work and conducting data management, which we gratefully acknowledged. This
study was funded by the Government of the Republic of Indonesia Department of Research
Technology and Higher Education, PTUPT-2018, under the contract 1859/UN1/DITLIT/DITLIT/LT/2018
NOMENCLATURE
vα = angle of attacks (AoA/deg)
ρ = density (kg/m3)
ω = vorticity tensor
bc = canard chord wide
crc = canard chord at canard center
h = aircraft hight
x/bc
= vortex centre height
xc/crc
= center distance concerning a canard
y/bc
= spanwise location
+
y = dimensionless wall distance
BBTF
= Bird's body type fuselage
CL = lift coefficient
Cp = pressure coefficient
HAC
= high AoA capability
M = Mach number
P = total pressure loss (Pa)
Q = Q-criterion: isosurface of instantaneous (s-2)
Remac
= mean aerodynamic chord Reynold number
RuV
= rolled-up vortex
S = rate-of-strain tensor
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314
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Sutrisno, Setyawan Bekti Wibowo, Sigit Iswahyudi and Tri Agung Rohmat
SPD
= surface pressure distribution
SBTF
= straight body type fuselage
Uc = axial canard vortex centre velocity (m/s)
U∞ = free stream velocity (m/s)
VBD
= vortex breakdown location
VCTH
= vortex core trajectory heights
Ws = wall-shear streamline (Pa)
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