JOURNAL OF AIRCRAFT
Vol. 53, No. 4, July-August 2016
Aerodynamic/Sonic Boom Performance Evaluation of Innovative
Supersonic Transport Configurations
Wataru Yamazaki∗
Nagaoka University of Technology, Nagaoka, Niigata 940-2188, Japan
and
Kazuhiro Kusunose†
Japan Aerospace Exploration Agency, Tokyo 182-8522, Japan
Downloaded by UNIVERSITY OF FLORIDA on March 17, 2017 | http://arc.aiaa.org | DOI: 10.2514/1.C033417
DOI: 10.2514/1.C033417
In this research, biplane wing/twin-body fuselage innovative configurations are discussed as a next-generation
supersonic transport candidate. Those aerodynamic performance as well as sonic boom performance are investigated
by using numerical approaches. The aerodynamic performance is evaluated by inviscid compressible Euler equations
using an unstructured mesh finite-volume method, and then the sonic boom performance is evaluated by an
augmented Burgers equation. Thanks to the successful interactions of shock waves between the biplane wing as well as
the twin-body fuselages, remarkable drag reduction and better sonic boom performance have been realized
simultaneously at our design Mach number of 1.7. The superiority of the proposed supersonic transport configuration
over conventional configurations is clearly demonstrated.
I.
wing configurations, which were 5 and 10% diamond-wedge wings
and Busemann’s biplane wing, were compared in the previous study.
The effectiveness of the twin-body fuselage concept with the all-wing
configurations has been demonstrated in the previous study. In the
present paper, the Licher-type biplane wing configuration, which is
known to be a better biplane wing configuration than the Busemann’s
biplane at lifted conditions [1,9], is also examined. The wave drag
characteristics as well as the sonic boom characteristics of the innovative SST configurations are investigated in this research by using
computational fluid dynamics (CFD) approaches. The freestream
Mach number in this study is set to our design Mach number of 1.7.
This paper is organized as follows. The CFD methodologies used in
this research to evaluate aerodynamic/sonic boom performance are
concisely described in Sec. II. The validity of the CFD methodologies
is discussed in Sec. III by comparing with experimental sonic boom
data. In Sec. IV, our innovative SST configurations are introduced,
and then those aerodynamic performances are discussed. In Sec. V,
those sonic boom performances are discussed. Finally, we provide
concluding remarks in Sec. VI.
Introduction
A
FUNDAMENTAL problem preventing large commercial
aircraft from supersonic flight is the creation of strong shock
waves, the effects of which are felt at the ground in the form of sonic
booms. Since the poor fuel efficiency as well as the sonic boom noise
at supersonic flight are mainly due to the effect of strong shock
waves, the reduction of shock strength is very important to realize
low-drag/low-boom supersonic transport (SST) configurations. A
reduction method of the shock strength due to the lifted wing has been
discussed in the literature [1–4] by introducing a supersonic biplane
wing concept. In this concept, the strength of the wave drag has been
successfully reduced by the interference of shock waves between the
biplane. According to Ref. [1], the wave drag at zero lift of the biplane
airfoil was reduced by nearly 90% compared to an equal-volume
diamond-wedge airfoil in two-dimensional inviscid simulations.
Inspired by the supersonic biplane wing concept, a twin-body
fuselage concept has also been proposed by the present authors for
more advanced SST configurations [5]. The same kind of twinfuselage concept was investigated in [6], in which fundamental wind
tunnel experiments have been performed at a Mach number of 2.7.
The main purpose of our concept of the twin-body fuselage is to
reduce the wave drag due to the volume of the aircraft’s fuselage.
Furthermore, unnecessary wave interactions between the fuselages
were minimized by using a shape optimization approach. According
to [5], over 20% total drag reduction was achieved by the optimized
twin-body fuselage compared to the Sears–Haack (SH) single-body
fuselage under the constraint of a fixed fuselage volume. The SH
body [7] is well known as the supersonic single-body configuration
that has the lowest wave drag for a specified volume and length.
The fusion of the two advanced concepts yields an innovative SST
configuration, which is a biplane wing/twin-body fuselage configuration. In [8], the innovative SST configuration was proposed, and
then its fundamental aerodynamic features were examined. Three
II.
Computational Methodologies
In this section, the computational methodologies used in this
research are concisely introduced.
A. Inviscid Flow Analysis
Three-dimensional supersonic inviscid flows are analyzed by an
unstructured mesh CFD solver of the TAS code [10,11]. Compressible Euler equations are solved by a finite-volume cell-vertex scheme.
The numerical flux normal to the control volume boundary is computed using the approximate Riemann solver of Harten–Lax–van
Leer–Einfelds–Wada [12]. The second-order spatial accuracy is
achieved by the unstructured monotonic upstream-centered scheme for
conservation laws approach [11,13] with Venkatakrishnan’s limiter
[14]. The implicit lower/upper symmetric Gauss–Seidel method for
unstructured meshes [15] is used for the time integration. Threedimensional unstructured meshes are generated using the TAS mesh
package, which includes surface mesh generation by an advancing
front approach [16,17] and tetrahedral volume mesh generation by a
Delaunay approach [18]. The high accuracy of this unstructured mesh
CFD approach has already been confirmed in the literature [11,19].
Received 26 February 2015; revision received 11 May 2015; accepted for
publication 14 September 2015; published online 22 February 2016.
Copyright © 2015 by Wataru Yamazaki and Kazuhiro Kusunose. Published
by the American Institute of Aeronautics and Astronautics, Inc., with
permission. Copies of this paper may be made for personal and internal use, on
condition that the copier pay the per-copy fee to the Copyright Clearance
Center (CCC). All requests for copying and permission to reprint should be
submitted to CCC at www.copyright.com; employ the ISSN 0021-8669
(print) or 1533-3868 (online) to initiate your request.
*Associate Professor.
†
Visiting Research Scientist.
B. Skin Friction Drag Estimation
In this research, skin friction drags of various SST configurations
are estimated by introducing simple algebraic skin friction models.
942
943
Downloaded by UNIVERSITY OF FLORIDA on March 17, 2017 | http://arc.aiaa.org | DOI: 10.2514/1.C033417
YAMAZAKI AND KUSUNOSE
Fig. 1 CFD evaluations of D-SEND#1 models.
Assuming that the boundary layer along the body is fully turbulent,
the skin friction drag coefficient can be estimated as
CDf Cf
Swet
Sref
(1)
where Cf is the averaged turbulent skin friction coefficient on the
wetted area of the body and Swet and Sref are, respectively, the wetted
area of the body and reference area. The skin friction coefficient for
turbulent boundary-layer conditions can be calculated by the
Prandtl–Schlichting flat-plate skin friction formula [20,21],
Cf 0.455
log10 Re2.58 1 0.144M2∞ 0.65
(2)
where Re and M∞ are, respectively, the Reynolds number and
freestream Mach number. In this research, the Reynolds number is
given at the cruise condition of Concorde (total length l of 62 m). Since
the speed of sound a∞ and kinematic viscosity ν∞ at the altitude of
18,000 m are, respectively, 295.069 m∕s and 1.1686 × 10−4 m2 ∕s
according to Ref. [22], the Reynolds number for the fuselage body is
calculated as
Re M∞ a∞ l
≅ 266 × 106
ν∞
(3)
The Reynolds number for the main wing is calculated in the same
manner with its mean chord length. The skin friction drag coefficients
of the fuselage and wing are separately estimated with the corresponding Reynolds numbers by using Eqs. (1) and (2). In [1,2,23], the
predicted friction drag based on the algebraic skin friction models is
compared to those based on viscous CFD computations. It has been
concluded that the simple algebraic skin friction models are reasonably
accurate for the prediction of friction drag in supersonic flows at
general angles of attack. We have to note that the simple models do not
take account of the effects of flow separations and reattachments at
large angles of attack.
C. Sonic Boom Analysis
The sonic booms on the ground are predicted by a nonlinear acoustic propagation solver Xnoise [24,25], which has been developed by
Japan Aerospace Exploration Agency (JAXA). An augmented Burgers equation is numerically solved by using an operator splitting
method, which takes into account the effects of nonlinearity, geometrical spreading, inhomogeneity of the atmosphere, thermoviscous
attenuation, and molecular vibration relaxation. In this approach,
initial (input) pressure distributions are extracted from CFD solutions
on the lower side of SST configurations (typically two fuselage
lengths below). Then, the propagation of the pressure distribution to
the ground is solved by the augmented Burgers equation.
III.
Validation of Computational Methodologies
The validity of the present computational methodologies is discussed in this section by comparing with the experimental data of the
D-SEND#1 project of JAXA‡ [26]. The sonic boom distributions
generated from an N-wave model (NWM) and a low-boom model
(LBM) at M∞ 1.58 and 1.59 are investigated. The entire lengths of
the NWM and LBM are, respectively, 5.6 and 8.0 m. In those
experiments, the sonic boom distributions were measured at about
3500 m away from the models.
The inviscid CFD analyses are performed at the freestream Mach
numbers to extract the initial pressure distributions as shown in Fig. 1.
In this research, the following volume grid generation procedure is
adopted for the accurate prediction of the initial pressure distributions,
which is schematically shown in Fig. 2. First, the tetrahedral elements are generated in a limited half-cylindrical region including the
semispan object as shown in Fig. 2a. Then, prismatic elements are built
up from the lateral surface of the half-cylinder covering a userspecified computational domain. The prismatic elements are generated
along the direction of the Mach angle as shown in Fig. 2b. The overall
views of the generated computational grid of the LBM are shown in
Figs. 2c and 2d. The inclination of the prismatic elements at the Mach
angle remarkably increases the accuracy of the initial pressure distribution. In Fig. 3, the pressure distributions extracted at two body
lengths below the LBM are compared between the inclination/
noninclination grids. Although the computational grid resolutions are
almost identical between them, much sharper pressure distribution is
obtained with the inclination of the prismatic elements at the Mach
angle. In Fig. 3, the extracted pressure distribution of the NWM, which
is also evaluated by a computational grid with the inclination of the
prismatic elements at the Mach angle, is also shown. The numbers of
computational grid points are about 2.6 million in both models.
Then, the sonic boom propagations are solved by the augmented
Burgers equation to compare to the experimental sonic boom distributions, the results of which are shown in Fig. 4. The fluctuations of the
experimental data in the range of 0.03 < t < 0.05 s are considered to
be the effect of atmospheric turbulence. Two major sonic boom
performance parameters of maximum pressure rise ΔP and impulse
Im are calculated from the final pressure distributions, and these
parameters are summarized in Table 1. Qualitative agreements with
the experimental data can be confirmed in both the NWM and LBM
cases, which indicates the validity of the present computational
methods.
‡
JAXA D-SEND Data available online at, http://d-send.jaxa.jp/d_send_e/
index.html [retrieved 30 January 2015].
Downloaded by UNIVERSITY OF FLORIDA on March 17, 2017 | http://arc.aiaa.org | DOI: 10.2514/1.C033417
944
YAMAZAKI AND KUSUNOSE
Fig. 2
IV.
Procedure of the computational grid generation for accurate sonic boom prediction.
Aerodynamic Performance of Wing/Body
Configurations
In this section, the aerodynamic performance of biplane wing/
twin-body fuselage SST configurations is discussed.
A. Definitions of Wing/Body Configurations
Unswept tapered biplane wing configurations are designed in this
research. The aspect and taper ratios are, respectively, set to about 7
and 0.25. The chord length of the main wing at the root section is
about l∕6. A vertical wing-tip plate is arranged between the upper/
lower wings to increase the two dimensionality of the flow around the
biplane wing. For appropriate shock interactions, the vertical distance
between the wings is shortened at the outer wing by adding a dihedral
angle to the lower wing. The Busemann-type and Licher-type biplane
wing configurations are discussed in this research. The section airfoil
thickness ratios of the Busemann-type biplane are set to 5% of the
chord length in both the upper and lower wings. The thickness ratios
of the Licher-type biplane (different between the upper/lower
wings) are determined to have the same sectional area with the
Busemann-type biplane wing. The mounting angle (β in Fig. 5) of the
Licher-type wing is set to 2 deg. The outlines of the biplane wings are
shown in Fig. 5. For comparison purposes, an unswept tapered
conventional wing configuration is also designed, the section shape
of which is a diamond-wedge airfoil. The thickness ratio is set to
10%, which has the same volume as the biplane wing configurations.
The outline of this wing is also included in Fig. 5.
With respect to the fuselage, three fuselage body configurations,
which have the same total volume of the fuselage, are investigated in
this research. A nonaxisymmetrical twin-body fuselage configura-
Fig. 3 Extracted initial pressure distributions of D-SEND#1 models.
945
YAMAZAKI AND KUSUNOSE
Downloaded by UNIVERSITY OF FLORIDA on March 17, 2017 | http://arc.aiaa.org | DOI: 10.2514/1.C033417
Fig. 5
Fig. 4
Outlines of wing configurations.
Comparison with experimental data of D-SEND#1 project.
tion has been designed by the present authors [5], minimizing the
inviscid drag coefficient at M∞ 1.7 and an angle of attack of 2 deg.
Ten design variables were used to represent various nonaxisymmetrical twin-body configurations, while the distance between the
twin bodies was fixed to l∕4. In this design process, an efficient
global design optimization approach via a Kriging response surface
model [27,28] was used. The expected improvement (EI) value [27]
can be evaluated with the Kriging model on the design variables
space, which indicates the potential for improvement considering
both the estimated function value and uncertainty in the response
surface model. Since the computational cost to evaluate the EI is
very small, the massive exploration of the design variables space
(maximization of the EI value) can be performed inexpensively with
a standard genetic algorithm. According to [5], a 23% total drag
reduction has been realized in the nonaxisymmetrical twin-body
optimal configuration compared to the single-body SH configuration
that has the same volume of the fuselage. This drag reduction has
been realized by successful wave interactions between the twin
bodies. For comparison purposes, single/twin-body SH configurations are also investigated in this research. The outlines of the three
fuselage configurations are shown in Fig. 6.
These wings and fuselages are merged to make SST wing/body
configurations. The following six combinations are investigated in
this research: 1) SH single-body with diamond-wedge/Busemann biplane wings, 2) SH twin body with diamond-wedge/Busemann
biplane wings, 3) nonaxisymmetrical twin body with Busemann/
Licher biplane wings.
The computational grids around the SST configurations are generated in the same manner as the previous section. The numbers of
computational grid points are about 3 million in all the configurations. The unstructured grid around a twin-body/biplane wing
configuration is visualized in Fig. 7.
B. Results and Discussion
In Figs. 8 and 9, the pressure contours around the SST configurations are shown at the condition of M∞ 1.7 and lift coefficient
CL 0.15. We can observe (successful) shock interactions between
the wings as well as the twin bodies. The inviscid drag polar curves of
these configurations at M∞ 1.7 are shown in Fig. 10. The same
Table 1
Model
NWM
LBM
Sonic boom performance of D-SEND#1 models
Experiment
ΔP, Pa Im, Pa · s
33.0
0.268
17.7
0.182
Simulation
ΔP, Pa
Im, Pa · s
30.7 (−7%)
0.258 (−4%)
15.0 (−16%) 0.168 (−7%)
Fig. 6
Outlines of fuselage body configurations.
Fig. 7 Unstructured grid around a SST configuration; left: overall view,
right: enlarged view.
reference area is used in all the cases for the evaluation of lift/drag
coefficients, which is the projected area of the main wing (the same in
all the wing configurations). The general range of the angle of attack
is from 0 to 5 deg, while that of the Licher biplane wing configuration
is from −2 to 3 deg due to the mounting angle of 2 deg. It can be
understood that the biplane wing concept and the twin-body fuselage
concept have individual drag reduction effects. We can summarize
that about 150 cts drag reduction is achieved by the adoption of
the biplane wing configuration, and then, separately, about another
60 ∼ 80 cts drag reduction is achieved by the adoption of the twinbody fuselage configuration. The best aerodynamic performance is
achieved by the nonaxisymmetrical twin-body fuselage configurations. In detail, the Licher biplane wing configuration has better aerodynamic performance than the Busemann biplane wing configuration
at the conditions of CL > 0.1.
The skin friction drag coefficients of the SST configurations are
estimated by the algebraic model of Eqs. (1) and (2). The total
(inviscid plus skin friction) drag polar curves of these configurations
at M∞ 1.7 are shown in Fig. 11. Although the algebraic model
does not take account of the effects of flow separations and re-
Downloaded by UNIVERSITY OF FLORIDA on March 17, 2017 | http://arc.aiaa.org | DOI: 10.2514/1.C033417
946
YAMAZAKI AND KUSUNOSE
Fig. 8
Pressure contours around SST configurations.
attachments, those effects are considered to be insignificant in the
considered range of the angle of attack. Since the twin-body/biplane
wing configurations have larger wetted areas than the single-body/
diamond-wedge wing configurations, the drag reduction effect
shortens by taking into account the viscous contributions. But the
superiorities of the twin-body/biplane wing configurations can be
still observed. The aerodynamic performance at M∞ 1.7 and
CL 0.15 is summarized in Table 2. Although all the configurations
have the same volume of the fuselage/wing, the nonaxisymmetrical
twin-body/Licher biplane wing configuration achieves a 38% total
drag reduction compared to the SH single-body/diamond-wedge
wing conventional configuration. It is, therefore, confirmed that the
proposed twin-body/biplane wing configuration is a promising candidate for the next-generation large-sized SST. Readers may notice
the low lift over drag ratio (L∕D) value of the proposed SST
configuration, which is less than 5, while that of Concorde is known to
be about 7 ∼ 8 (at M∞ 2.0). This was also discussed in [8], in which
it was concluded that the major reason is its larger fuselage volume
than that of Concorde. Actually, the fuselage volume of our proposed
configuration is approximately four times more than that of Concorde,
which yields the lower L∕D value than Concorde in this study.
Although the aerodynamic effectiveness of the proposed configuration
has been confirmed in this research, further sophistication of the
configuration, such as aerodynamic shape optimization of the biplane
wing and resizing of the fuselage volume, is essential for the proposition of a commercially viable SST configuration.
V.
Sonic Boom Performance of Wing/Body
Configurations
The sonic boom performance of the proposed SST configurations
is also investigated at M∞ 1.7 and CL 0.15. In this study, the
pressure distributions are extracted at two fuselage lengths below
the SST configurations on the symmetry plane as is shown in Fig. 12.
In the sonic boom propagation analyses, standard atmosphere
temperature/humidity profiles are used. The fuselage length and the
cruise altitude are, respectively, set to 62 and 18,000 m and are given
from the conditions of Concorde.
The pressure distributions extracted below the SST (on the line
shown in Fig. 12) are compared in Fig. 13. Larger pressure variations
can be observed in the diamond-wedge wing configurations at
the medium of the distributions. The propagation of the pressure
distribution to the ground is solved by the augmented Burgers
equation, and then the pressure distributions on the ground are
compared in Fig. 14. In Figs. 13 and 14, the distributions of “SH
Single + Diamond Wing” and “SH Twin + Diamond Wing” are
almost equivalent. In addition, the distributions of “SH Single +
Busemann”, “SH Twin + Busemann,” and “Nonaxisymmetric Twin +
Busemann” are almost equivalent. These results indicate that the
sonic boom characteristics can be primarily classified according to
the wing configurations in this study. The diamond-wedge wing
configurations have typical N-wave distributions, while the other
configurations have two peaks in the distributions. In other words, the
insignificance of the fuselage body configurations to the sonic boom
947
Downloaded by UNIVERSITY OF FLORIDA on March 17, 2017 | http://arc.aiaa.org | DOI: 10.2514/1.C033417
YAMAZAKI AND KUSUNOSE
Fig. 9 Pressure contours at the 60% semispan section of SST configurations.
Fig. 10 Inviscid drag polar at M∞ 1.7.
performance can be confirmed in this study. Four major sonic boom
performance parameters of ΔP, Im, the sound pressure level Lp , and
the A-weighted sound exposure level LAE are calculated from the
pressure distributions on the ground, and these parameters are
summarized in Table 3. Lp is calculated from the following
formulation:
Lp 20 log10 ΔP∕20 × 10−6 (4)
LAE is evaluated by the software of BoomMetre [29], which was
developed by JAXA. The best sonic boom performance is obtained
Table 2 Aerodynamic performance of SST configurations at
M∞ 1.7 and CL 0.15 (Drag coefficients in drag count, 1 drag
count 0.0001)
Configuration
SH single + diamond
SH single + Busemann
SH twin + diamond
SH twin + Busemann
Nonaxisymmetric twin + Busemann
Nonaxisymmetric twin + Licher
α,
deg
2.66
2.35
2.57
2.47
2.50
0.16
Inviscid Friction
drag
drag
511.8
74.8
328.4
112.1
428.4
89.1
276.7
128.1
245.6
128.6
234.8
128.4
Total
drag
586.7
440.5
517.5
404.8
374.2
363.2
Downloaded by UNIVERSITY OF FLORIDA on March 17, 2017 | http://arc.aiaa.org | DOI: 10.2514/1.C033417
948
YAMAZAKI AND KUSUNOSE
Fig. 11 Total drag polar at M∞ 1.7.
Fig. 14 Final pressure distributions at the ground.
Fig. 12 Extraction of pressure distribution for sonic boom analysis.
functions are indicated. Although LAE is known as a better metric
to express the subjective loudness of sonic booms than ΔP [30],
we can observe a positive correlation between ΔP and LAE in the
present case. We can also observe positive correlations between ΔP
and Im, CD and Im, and CD and LAE . The sound pressure level
of the nonaxisymmetrical twin-body/Licher biplane wing configuration is not much more than 124 dB sound pressure level (SPL),
while that of Concorde is known to be over 130 dB SPL. According to
[31], the target value of ΔP for next-generation low-boom SST is
considered to be about 24 Pa, i.e., 121.6 dB SPL. For the proposition
of a more favorable next-generation SST configuration, therefore, not
only the shock wave interactions between the bodies and between
Table 3
Fig. 13 Initial pressure distributions at extraction line.
in the nonaxisymmetrical twin-body fuselage with the Licher
biplane wing, the maximum pressure rise of which is 38% lower than
the single-body/diamond-wedge wing configuration. In Fig. 15,
the relationships between aerodynamic/sonic boom performance
Sonic boom performance of SST configurations at
M∞ 1.7 and CL 0.15
Configuration
SH single + diamond
SH single + Busemann
SH twin + diamond
SH twin + Busemann
Nonaxisymmetric twin + Busemann
Nonaxisymmetric twin + Licher
Im,
ΔP,
LP ,
dB SPL Pa · s
Pa
49.19 127.82 5.15
32.72 124.28 3.62
48.58 127.71 5.04
34.40 124.71 3.75
33.18 124.40 3.61
30.59 123.69 2.95
LAE ,
dB
84.69
81.46
84.53
81.82
81.47
80.29
949
Downloaded by UNIVERSITY OF FLORIDA on March 17, 2017 | http://arc.aiaa.org | DOI: 10.2514/1.C033417
YAMAZAKI AND KUSUNOSE
Fig. 15 Relationships between aerodynamic/sonic boom performance functions.
wings but also its detailed shape optimization minimizing sonic
boom will be required.
VI.
Conclusions
In this research, twin-body fuselage/biplane wing configurations
for a large-sized supersonic transport (SST) have been discussed. The
wave drag characteristics/reduction by these innovative concepts
have been discussed at a design freestream Mach number of 1.7 by
using inviscid computational fluid dynamics (CFD) computations.
The increment of skin friction drag has also been discussed, using
the standard algebraic (turbulent) skin friction models based on
the wetted areas of SST configurations. Furthermore, the sonic
boom performance of the innovative SST configurations has also
been discussed. As one of the twin-body fuselage configurations, the
nonaxisymmetrical twin-body configuration that the authors had
designed beforehand was also investigated. As biplane wing configurations, Busemen-/Licher-type biplanes were investigated.
The biplane wing concept and the twin-body fuselage concept
have individual drag reduction effects. The combination of the
nonaxisymmetrical twin-body fuselage with the Licher biplane wing
achieved the best aerodynamic performance among the investigated
cases. The proposed SST configuration achieved a 38% total drag
reduction compared to the single-body/diamond-wedge wing conventional configuration under constraints of a fixed volume and
fixed length of the fuselage. This result indicates the remarkable
aerodynamic superiority of the proposed twin-body/biplane wing
configuration.
The sonic boom characteristics of the SST configurations were
primarily classified according to the wing configurations. The sonic
boom performance parameters became better by the adoption of the
biplane wing configurations. On the other hand, the fuselage body
configurations had less effect on the sonic boom performance in this
study. The best sonic boom performance was also obtained in the
nonaxisymmetrical twin-body fuselage with the Licher biplane wing,
of which the maximum pressure rise at the ground was 38% lower
than the single-body/diamond-wedge wing conventional configuration.
For a more sophisticated SST configuration, aerodynamic shape
optimization of the biplane wing and the resizing of the fuselage
volume are essential. The design (optimization) of a wing/body
fairing should also be considered, and the supersonic drag may be
reduced more. Furthermore, low-boom design optimization will
be also required for the proposition of more realistic low-boom/
low-drag SST configurations. Since the fundamental availability of
the proposed concept has been confirmed from the viewpoints of
aerodynamics/sonic boom acoustics in this research, the authors will
continue further development research studies to demonstrate its
inclusive availability by additional multidisciplinary investigations.
Acknowledgments
The authors are very grateful to Masashi Kanamori, Yusuke Naka,
Yoshikazu Makino, and Keiichi Murakami of Japan Aerospace
Exploration Agency for providing them with their nonlinear acoustic
propagation solver Xnoise as well as for their helpful advice. The
authors also would like to thank Naohiko Ban of Nagaoka University
of Technology for his help in making CAD models of supersonic
transport configurations.
References
[1] Kusunose, K., Matsushima, K., and Maruyama, D., “Supersonic
Biplane-A Review,” Progress in Aerospace Sciences, Vol. 52, No. 6,
2011, pp. 53–87.
doi:10.1016/j.paerosci.2010.09.003
[2] Kusunose, K., Matsushima, K., Obayashi, S., Furukawa, T., Kuratani,
N., Goto, Y., Maruyama, D., Yamashita, H., and Yonezawa, M., “Aerodynamic Design of Supersonic Biplane: Cutting Edge and Related
Topics,” The 21st Century COE Program International COE of Flow
Dynamics Lecture Series, Vol. 5, Tohoku Univ. Press, Sendai, Japan,
2007.
[3] Matsushima, K., Kusunose, K., Maruyama, D., and Matsuzawa, T.,
“Numerical Design and Assessment of a Biplane as Future Supersonic
Transport,” Proceedings of the 25th ICAS Congress, International
950
[4]
[5]
[6]
[7]
Downloaded by UNIVERSITY OF FLORIDA on March 17, 2017 | http://arc.aiaa.org | DOI: 10.2514/1.C033417
[8]
[9]
[10]
[11]
[12]
[13]
[14]
[15]
[16]
[17]
YAMAZAKI AND KUSUNOSE
Council of the Aeronautical Sciences Paper 2006-3.7.1, Bonn,
Germany, 2006.
Matsushima, K., Maruyama, D., Kusunose, K., and Noguchi, R.,
“Extension of Busemann Biplane Theory to Three Dimensional Wing
Fuselage Configurations,” Proceedings of the 27th ICAS Congress,
International Council of the Aeronautical Sciences Paper 2010-2.8.1,
Bonn, Germany, 2010.
Yamazaki, W., and Kusunose, K., “Aerodynamic Study of Twin-Body
Fuselage Configuration for Supersonic Transport,” Transactions of the
Japan Society for Aeronautical and Space Sciences, Vol. 56, No. 4,
2013, pp. 229–236.
doi:10.2322/tjsass.56.229
Wood, R. M., Miller, D. S., and Brentner, K. S., “Theoretical and
Experimental Investigation of Supersonic Aerodynamic Characteristics
of a Twin-Fuselage Concept,” NASA TP-2184, 1983.
Sears, W. R., “On Projectiles of Minimum Wave Drag,” Quarterly of
Applied Mathematics, Vol. 4, No. 4, 1947, pp. 361–366.
Yamazaki, W., and Kusunose, K., “Biplane Wing/Twin-Body-Fuselage
Configuration for Innovative Supersonic Transport,” Journal of Aircraft, Vol. 51, No. 6, 2014, pp. 1942–1952.
doi:10.2514/1.C032581
Licher, R. M., “Optimum Two-Dimensional Multiplanes in Supersonic
Flow,” Douglass Aircraft Co. Rept. SM-18688, Santa Monica, CA,
1955.
Nakahashi, K., Ito, Y., and Togashi, F., “Some Challenges of Realistic
Flow Simulations by Unstructured Grid CFD,” International Journal
for Numerical Methods in Fluids, Vol. 43, Nos. 6–7, 2003, pp. 769–783.
doi:10.1002/(ISSN)1097-0363
Yamazaki, W., Matsushima, K., and Nakahashi, K., “Drag Prediction,
Decomposition and Visualization in Unstructured Mesh CFD Solver of
TAS-Code,” International Journal for Numerical Methods in Fluids,
Vol. 57, No. 4, 2008, pp. 417–436.
doi:10.1002/(ISSN)1097-0363
Obayashi, S., and Guruswamy, G. P., “Convergence Acceleration of a
Navier–Stokes Solver for Efficient Static Aeroelastic Computations,”
AIAA Journal, Vol. 33, No. 6, 1995, pp. 1134–1141.
doi:10.2514/3.12533
Burg, C. O. E., “Higher Order Variable Extrapolation for Unstructured
Finite Volume RANS Flow Solvers,” AIAA Paper 2005-4999, 2005.
Venkatakrishnan, V., “On the Accuracy of Limiters and Convergence to
Steady State Solutions,” AIAA Paper 1993-0880, 1993.
Sharov, D., and Nakahashi, K., “Reordering of Hybrid Unstructured
Grids for Lower-Upper Symmetric Gauss-Seidel Computations,” AIAA
Journal, Vol. 36, No. 3, 1998, pp. 484–486.
doi:10.2514/2.392
Ito, Y., and Nakahashi, K., “Direct Surface Triangulation Using
Stereolithography Data,” AIAA Journal, Vol. 40, No. 3, 2002, pp. 490–
496.
doi:10.2514/2.1672
Ito, Y., and Nakahashi, K., “Surface Triangulation for Polygonal Models
Based on CAD Data,” International Journal for Numerical Methods in
[18]
[19]
[20]
[21]
[22]
[23]
[24]
[25]
[26]
[27]
[28]
[29]
[30]
[31]
Fluids, Vol. 39, No. 6, 2002, pp. 75–96.
doi:10.1002/(ISSN)1097-0363
Sharov, D., and Nakahashi, K., “Hybrid Prismatic/Tetrahedral Grid
Generation for Viscous Flow Applications,” AIAA Journal, Vol. 36,
No. 2, 1998, pp. 157–162.
doi:10.2514/2.7497
Umeda, M., Takenaka, K., Hatanaka, K., Hirayama, D., Yamazaki, W.,
Matsushima, K., and Nakahashi, K., “Validation of CFD Capability for
Supersonic Transport Analysis Using NEXST-1 Flight Test Results,”
AIAA Paper 2007-4437, 2007.
Raymer, D. P., Aircraft Design: A Conceptual Approach, 4th ed., AIAA
Education Series, AIAA, Reston, VA, 2006, pp. 327–330.
Krasnov, N. F., Aerodynamics of Bodies of Revolution, edited by Morris,
D. N., Elsevier, New York, 1970, pp. 379–385.
The Japan Society for Aeronautical and Space Sciences, Handbook of
Aerospace Engineering, 3rd ed., Japan Soc. for Aeronautical and Space
Sciences, Maruzen, Tokyo, 2005, pp. 8–12 (in Japanese).
Hu, R., “Supersonic Biplane Design via Adjoint Method,” Ph.D. Thesis,
Dept. of Aeronautics and Astronautics, Stanford Univ., Stanford, CA,
2009.
Yamamoto, M., Hashimoto, A., Takahashi, T., Kamamura, T., and
Sakai, T., “Long-Range Sonic Boom Prediction Considering Atmospheric Effects,” Proceedings of Inter Noise 2011 [CD-ROM], Curran
Associates, Inc., Red Hook, NY, 2011.
Yamamoto, M., Hashimoto, A., Takahashi, T., Kamakura, T., and Sakai,
T., “Numerical Simulations for Sonic Boom Propagation Through an
Inhomogeneous Atmosphere with Winds,” AIP Conference Proceedings, Nonlinear Acoustics, American Inst. of Physics, College Park,
MD, 2012, pp. 339–342.
Naka, Y., “Sonic Boom Data from D-SEND#1,” JAXA RM-11-010E,
Tokyo, 2012.
Jones, D. R., Schonlau, M., and Welch, W. J., “Efficient Global
Optimization of Expensive Black-Box Functions,” Journal of Global
Optimization, Vol. 13, No. 4, 1998, pp. 455–492.
doi:10.1023/A:1008306431147.
Yamazaki, W., and Mavriplis, D. J., “Derivative-Enhanced Variable
Fidelity Surrogate Modeling for Aerodynamic Functions,” AIAA Journal, Vol. 51, No. 6, 2013, pp. 126–137.
doi:10.2514/1.J051633
Makino, Y., Naka, Y., Hashimoto, A., Kanamori, M., Murakami, K., and
Aoyama, T., “Sonic Boom Prediction Tool Development at JAXA,”
Journal of the Japan Society for Aeronautical and Space Sciences,
Vol. 61, No. 7, 2013, pp. 237–242 (in Japanese).
Leatherwood, J. D., and Sullivan, B. M., “Laboratory. Study of Effects
of Sonic Boom Shaping on Subjective Loudness and Acceptability,”
NASA TP 3269, 1992.
Yoshida, K., “Low Sonic Boom Design Technology for Next Generation
Supersonic Transport,” Journal of the Acoustical Society of Japan,
Vol. 67, No. 6, 2011, pp. 245–250 (in Japanese).