Aerodynamic Design of Unmanned and Scaled Supersonic ...

European Congress on Computational Methods in Applied Sciences and Engineering
ECCOMAS 2004
P. Neittaanmäki, T. Rossi, S. Korotov, E. Oñate, J. Périaux, and D. Knörzer (eds.)
Jyväskylä, 24—28 July 2004
AERODYNAMIC DESIGN OF UNMANNED AND SCALED
SUPERSONIC EXPERIMENTAL AIRPLANE IN JAPAN
Kenji Yoshida * and Yoshikazu Makino *
* Institute of Space Technology and Aerodynamics
Japan Aerospace Exploration Agency (JAXA), Tokyo, Japan
e-mails: yoshida.kenji@jaxa.jp, yoshikazu.makino@jaxa.jp
Key words: Drag Reduction, Arrow Wing, Natural Laminar Flow, Inverse Design Method,
Adjoint Method
Abstract. This paper describes the aerodynamic design of the unmanned and scaled
supersonic experimental airplanes promoted by Japan Aerospace Exploration Agency (JAXA).
The main goal of the experimental airplane program is to develop an advanced aerodynamic
design technology for the next generation SST. In the program conventional and innovative
drag reduction technologies were mainly considered. The first airplane had no propulsion
system and was to be launched by a solid rocket booster. In the aerodynamic design, we
applied an optimum combination of classical pressure drag reduction concepts and a new
friction drag reduction concept. The classical concepts are grounded in supersonic linear
theory and involve the application of a suitable arrow planform, a warped wing and an arearuled
body. The new concept is an original technical challenge involving the design of a
natural laminar flow wing with a subsonic leading edge at supersonic speed. This concept
consists of both an ideal pressure distribution to delay transition and an original CFD-based
inverse design method. The validity of the concept has already been confirmed qualitatively in
the wind tunnel tests. However, we have never validated it quantitatively, because of the
inherent freestream turbulence that exists in supersonic wind tunnels. Therefore, the main
objective of the flight test is to validate the NLF wing concept. The second airplane has two
large jet engines under the wing. In the aerodynamic design, an original CFD-based optimum
design method was developed to reduce the interference drag between the airframe and the
engine nacelles. Using this method, we designed a non-axisymmetrical area-ruled body and
an optimum nacelle shape for the second experimental airplane. JAXA’s CFD calculations
confirmed that the strong interference drag due to large nacelles was remarkably reduced
over the whole Mach number range. The main objective of the flight test is mainly to validate
the non-axisymmetrical area-ruled body concept.
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K. Yoshida and Y. Makino
1 INTRODUCTION
After the development of the Concord, the U.S.A. and Europe investigated several
advanced technologies for the challenge of realizing the development of the next generation
SST. In general, the challenge can be undertaken within the framework of an international
collaboration. We at Japan have to advance our technology level in order to join the
collaboration and support the development.
Japan Aerospace Exploration Agency (JAXA) is promoting the National EXperimental
Supersonic Transport (NEXST) program. This program was initiated in 1997 and will last
until 2005. It consists of fundamental research activity and flight test program. The main
objectives of the fundamental research activity are to advance present technologies in Japan
and to create some innovative technologies for the next generation SST in the future. Since
Japan has high advanced capability in CFD technology, the principal goal is focused on
developing a CFD-based optimum aerodynamic design technology.
In the flight test program, we have designed and developed two kinds of unmanned and
scaled experimental airplanes. The main objective of the first airplane is mainly to validate
supersonic drag reduction technology by the flight test of a pure aerodynamic configuration
without any propulsion system. The main objective of the second airplane is to validate the
CFD-based optimum design technology by the flight test of a more complicated configuration
with two engines and nacelles under the wing.
The first airplane is called “NEXST-1”
airplane and will be launched by a solid rocket
booster (see Figure1). The aerodynamic design
of the airplane was undertaken in the following
two steps: i) applying the well-known clasical
drag reduction concepts for reducing pressure
drag, and ii) designing an original supersonic
natural laminar flow (NLF) wing concept for
reducing friction drag. The final aerodynamic
design and production of the NEXST-1 airplane
is already completed. The flight test plan is
summarized in Figure 1. The aerodynamic
forces, surface pressures and transition data will
Woomera
Separation
Launch
α-sweep at M=2 and low
18,000 m
15,000 m
be measured at two α sweep tests. In its production, the severe roughness criterion of
maximum height within 1 µ m was required for the successful transition measurement.
The second airplane is called “NEXST-2” and is a jet-powered vehicle. In the aerodynamic
design of the NEXST-2 airplane, first of all, the following basic design concepts were
applied: an arrow planform with slightly higher aspect ratio than the NEXST-1 airplane, a
modified warped wing with two embedded nacelles, an NLF wing at inboard, a supersonic
leading edge at outboard. Secondly, an original non-axisymmetrical area-ruled body with an
optimum nacelle configuration was designed using a CFD-based optimum design method
developed in JAXA. This concept was utilized in reducing the interference drag between the
airframe and two large nacelles. The basic aerodynamic design of the NEXST-2 airplane was
11,000 m
α-sweep at M=2 and high
Recovery
Figure 1. Flight test plan of NEXST-1 airplane
2

K. Yoshida and Y. Makino
completed in the middle of 2003. The aim of the
flight test will be to ensure the practicality of
applying the design method to real aircraft
design. The flight test plan of the NEXST-2
airplane is summarized in Figure 2. Several
technical data will be measured at supersonic
and subsonic flight.
The first flight test of the NEXST-1 airplane
was conducted on the 14 th of July in 2002.
Unfortunately the test was a failure, because of
unexpected early separation of the booster due
to an electric short on the firing system of
separation bolts. The failure resulted in the
M=0.8 at H=12 km
NEXST-2 program being frozen until the success of the flight test of the NEXST-1 airplane.
Since then, we have improved and redesigned the whole system. The next flight test is
scheduled for the end of 2004. We really do our best to realize the successful fight test of the
NEXST-1 airplane.
Present paper describes details of the aerodynamic design of both experimental airplanes.
2 AERODYNAMIC DESIGN OF NEXST-1 AIRPLANE
In general, we have the following standpoints in the aerodynamic design of the advanced
SST: reducing supersonic drag, reducing sonic boom, improving subsonic aerodynamic
performance, compromising aerodynamics and structures, and suppressing aerodynamic noise.
However as the first step of developing these advanced technologies, we mainly focused on
improving lift-to-drag ratio (L/D) at supersonic and subsonic speed in the flight test program,
because improving L/D is the most key technology in all development of SST. Furthermore,
reducing sonic boom and improving take-off/landing performance with high lift device are
also important. Therefore, they are investigated as fundamental reserach activities.
In the aerodynamic design of the NEXST-1 airplane, main target was placed on reducing
supersonic drag only, because this configuration is clean – that is, it is an aerodynamically
pure shape. Therefore, we can obtain the optimum effect of combining several drag reduction
concepts. In the NEXST program, we developed advanced aerodynamic design technology
according to the following philosophy: to design mathematically along the logical process
without any empirical parameters and to incorporate some innovative technologies.
To design the NEXST-1 airplane, first of all, we had to specify some typical requirements
on an expected real SST aircraft such as fuselage length, volume, tail shape and position. We
selected the following dimensions, referring the study [1] of JADC (Japan Aircraft
Development Corporation) and other foreign papers: M=2 (cruise Mach number), C L =0.1
(cruise lift coefficient), H=15km (flight altitude), S=9000 ft 2 (wing area), L=300 ft (fuselage
length), V=30000 ft 3 (fuselage volume for 300Pax) and scaled tail configurations of the
Concorde. Then, we selected a scale ratio of 11% of the real SST aircraft dimensions above.
However, tail cone length of the scaled fuselage was slightly extended to keep the
fire-on
take-off
M=1.7 at H=15 km
Test region
separation
data-link
landing
engine cut-off
X-38 scaled
model
Figure 2. Flight test plan of NEXST-2 airplane
3

K. Yoshida and Y. Makino
requirement of volume capacity for parachute system used in a recovery phase of the flight
test. Therefore, the tail shape of the scaled airplane is not similar as the real SST aircraft.
2.1 Design Concepts
Supersonic drag generally consists of pressure drag and friction drag. The pressure drag is
also divided into lift-dependent drag and wave drag due to volume. Furthermore, the liftdependent
drag is composed by vortex drag and wave drag due to lift. This drag breakdown is
based on supersonic linear theory [2].
We summarized our design concepts
incorporated in the aerodynamic design of the
NEXST-1 airplane in Figure 3. They are i)
Arrow Planform, ii) Warped Wing, iii) Arearuled
Body and iv) Supersonic NLF Wing. The
concepts of i), ii) and iii) are well known to be
effective in reducing lift-dependent drag by i),
ii) and wave drag due to volume by iii). They
are based on supersonic linear theory and their
design procedures were already derived in the
development of the Concorde. The concept of
iv) was originally developed in this program [3].
1. Arrow planform
・AR=2.2(S=10.12 m 2 )
2. Warped Wing
3. Area-ruled body
Design point : C L =0.1 @ M=2.0
11.5 m
4. Supersonic
NLF wing
Recovery system
Figure 3. Design concepts of NEXST-1
airplane
1) Arrow Planform Concept
In general, a higher aspect ratio is also aerodynamically desirable at supersonic speed.
Furthermore, there is an optimum slenderness ratio in supersonic linear theory [2]. The
slenderness ratio (s/l) of a wing planform is defined as a ratio of its semi-span length (s) to its
maximum streamwise length (l). However, the planform with such an optimum slenderness
ratio usually has a highly swept leading edge. A higher aspect ratio and a highly swept leading
edge generally lead to some structural penalties. Therefore, we need to compromise both the
increase of aspect ratio and the optimum slenderness ratio, to design a practical wing under
some structural constraints. Arrow planform with subsonic leading edge is well known to be
very effective in compromising. Therefore, the first principle for reducing supersonic drag is
to select an arrow planform with a suitable s/l near the optimum one taking account of
stractural constraints.
2) Warped Wing
To reduce the lift-dependent drag, one of the best ways is to adopt an ideal combination of
camber and twist distribution to realize an optimum load distribution. Such a cambered and
twisted surface is usually called “warped surface”. Carlson et al developed a numerical
method for estimating drag and designing a warped surface in 1974 [4]. They expressed an
optimum load as a combination of several prescribed elementary loads, and optimized their
combination coefficients by applying variational principle. As a result of the optimization,
they obtained an optimum camber and twist distribution to realize the lowest lift-dependent
4.72 m
4

K. Yoshida and Y. Makino
drag under the limitation of design condition and planform constraints.
The key point of warp design is to suppress theoretical infinite load at leading edge. This
usually leads to a certain leading edge droop to achieve an attached flow condition, because
leading edge separation vortex is induced by local high angle of attack due to highly swept
leading edge. This is the second principle for reducing supersonic drag.
3) Area-Ruled Body
According to the supersonic slender body theory [5], wave drag due to volume of a
complete aircraft is generally related to so-called supersonic cross sectional area distribution.
This area means the projection of oblique cross sectional area of airplane cut by Mach cone
on a plane vertical to streamwise direction. The optimum axisymmetric body with the lowest
wave drag due to volume was already derived using this formulation. It was called “Sears-
Haack body”. If a wing and tails of a complete aircraft are specified, we suppose that fuselage
geometry should be improved to adjust total supersonic area distribution to that of Sears-
Haack body. It is very effective to reduce the wave drag due to total volume of the aircraft.
This improved fuselage is generally called area-ruled body. This rule is the third principle of
reducing supersonic drag, especially interference drag between wings, tails and fuselage [5].
4) Natural Laminar Flow Wing
It is well expected that an aerodynamic optimum combination of the pressure drag
reduction concepts mentioned above has large effect in reducing supersonic cruise drag.
However, it is not easy to obtain the maximum gain of drag reduction effect, because we must
consider several constraints that are not included in linear theory. Therefore, any other drag
reduction concept is necessary to improve the L/D of a future advanced SST.
Reducing friction drag is one of the candidates. A laminar airfoil design concept is usually
based on suppressing Tollmien-Schlichting (T-S) wave instability. For a low aspect ratio wing
with highly swept leading edge, transition due to cross-flow (C-F) instability is dominant at
forward part of the wing. First of all, an optimum pressure distribution must be found for
suppressing the C-F instability. The key point is to reduce the region generating cross-flow.
Cross-flow is produced by chordwise pressure gradient. At the front part of the wing, there is
always severe acceleration. Therefore, it is very effective to set narrow acceleration region.
This leads to a pressure distribution with steep gradient at front.
Since the T-S instability becomes dominant after mid-chord of the wing, gradual
acceleration is effective to suppress the T-S instability. Fortunately, most of SST planforms
have supersonic trailing edge and require no recovery of trailing edge pressure. As a result, it
was found that pressure distribution of the NLF wing had to have steep acceleration at front
and gradual acceleration from the front to the trailing edge. At the first step of the NLF wing
design, we already found an optimum pressure distribution to delay transition [6], using a
practical transition prediction method (SALLY code [7]) based on e N method [8].
As the next step, we must solve a so-called inverse design problem to achieve this pressure
distribution. Therefore, we developed an original CFD-based inverse method [9]. This inverse
method consists of iteration loop for the following two routines: i) flow estimation using a
CFD code, and ii) modifying the geometry based on the difference between target and each
5

K. Yoshida and Y. Makino
estimated pressure distribution using supersonic linear theory. The supersonic NLF wing
design is the forth principle for reducing supersonic drag.
2.2 Design Process and Results
Our aerodynamic design process consisted of two phases [3]. At the first phase, we
designed a baseline configuration using a supersonic linear theory, namely lifting surface
theory and slender body theory. In this phase, the NLF wing design concept was not included.
At the second phase, we conducted some refinements of the baseline configuration using a
JAXA’s CFD (Navier-Stokes) code. Then we improved a lift-to-drag ratio (L/D) applying the
NLF wing concept. Design results at each process are described in the following sections.
1) 1 st Design Phase
(a) Planform design
In general, parameters characterizing arrow planform are wing area (S), aspect ratio (AR),
slenderness ratio (s/l), taper ratio (λ), leading edge sweep angle (Λ LE ), trailing edge sweep
angle (Λ TE ), and spanwise kink position (ε). In our planform study, some of those parameters
were determined referring to the typical arrow planform proposed by Douglas. It was shown
in Figure 4 as “H7-Baseline”. Among those parameters, we supposed major parameters were
AR, s/l, Λ LE,i (for inner wing), Λ TE,o (for outer wing) and ε T (for trailing edge). Here Λ TE =0°
was set for inner wing. The range of s/l was selected near the optimum value, which was
about 0.3 in this cruise Mach number. On the other hand, AR was taken from 1.8 to 2.2,
considering moderate balance between aerodynamics and structure.
First of all, according to these assumptions, ninety-nine planform candidates were designed
without any aerodynamic consideration. Second, the lift-dependent drag characteristics of
those planforms were estimated using a lifting surface theory [4] as a flat plate condition. One
of the criteria selecting an optimum planform was to pick up the planform with lower drag
characteristics than Douglas’s one. As another criterion, we chose the planform with lower
derivative of the drag to Mach number at design point. Finally eight planforms shown in
Figure 4 were selected for the next step.
(b) Warp design
Supersonic Lifting Surface Theory
NEXST-1 Warped Configuration: Geometry
Carlson's Optimum
(Carlson’s method)
For each eight planform, each warped surface
Lift-dependent drag
was designed, and each lift-dependent drag at
4
2
0
-2
-4
design point was estimated using Carlson's method
-6
-8
[4]. As shown in Figure 4, we finally selected the
0 50 100 150 x(ft
planform indicated as “H8-1st Baseline” with
AR=2.2 and Λ LE =66/61.2 degrees (inner/outer
wing).
However, in order to design a complete warped
wing, chordwise and spanwise thickness
distributions are necessary. The spanwise thickness Figure 4. Planform and Warp Design Results
ratio distribution was determined referring to
typical thickness ratio distribution of a foreign second generation SST configuration rather
8*z(ft)
6

K. Yoshida and Y. Makino
than that of Concorde. It has thickness ratio of 3.7% near wing root to keep enough space of
landing gear and 3% at outer wing to satisfy structural constraints [3]. The second generation
SST’s and Concorde’s spanwise thickness ratio distributions were shown in Figure 5. On the
other hand, we applied the chordwise thickness distribution of NACA 4-digit series, because
it has simple analytical formulation. Figure 4 also indicates each airfoil shape of the designed
warped wing.
(c) Area-ruled body design
Spanwise (t/c) max Distribution
0.06
Before applying area-ruled body design,
horizontal/vertical tail locations and wing location
0.05
were determined by referring to similar foreign
0.04
SST configurations. As an initial fuselage
0.03
configuration, we assumed a straight cylindrical
fuselage except nose and tail cones to have the
requirement of fuselage volume constraint.
According to the design procedure mentioned
above, a supersonic cross sectional area
distribution of the area-ruled body was estimated.
Figure 6 shows all components of the supersonic
area distributions. The area of Sears-Haack body
was estimated under the condition of the same
volume as the initial fuselage with wing and tails.
The area distribution of the area-ruled body was
calculated by subtracting the area of wing and
tails from the area of Sears-Haack body. Finally
the area-ruled body configuration was determined
from the estimated area distribution with
axisymmetrical body approximation. This wingbody-tails
configuration was called “1st
Configuration” [3].
(t/c)max
0.02
0.01
0
NEXST-1
Boeing 2nd Generation Type SST Type
Concorde Type
0 0.2 0.4 0.6 0.8 1 y/s
Figure 5. Spanwise thickness ratio distribution
Area(m 2 )
0.6
0.2
Supersonic Area Distribution: NEXST-1
M des =2.0
(t/c) Boeing Type
0.4
Sears-Haack Body
Area-Rule Body
Straight Body
Wing
V-Tail
H-Tail
0
0 5 10 x(m)
Figure 6. Supersonic cross sectional area
2) 2 nd Design Phase
(a) CFD analysis of 1st Configuration
At the first step of this phase, we analyzed the aerodynamic characteristics of the “1st
Configuration” using a CFD (Navier-Stokes) code [10] developed by JAXA under the full
turbulent condition. We found the following major differences [3, 11]: i) loss of total lift at
the same angle of attack, ii) increment of minimum drag, and iii) loss of lift at minimum drag
condition. The differences were mainly originated in the influence of wing thickness and body,
especially strong interference of an area-ruled body with static pressure on upper surface,
because they were not included in linear theory. The difference of (ii) and (iii) means the loss
of warp effect. The main reason was based on the load deficit near leading edge, compared
with the optimum load designed by Carlson method.
(b) Refinement by quasi-inverse design
To improve the drag characteristics of the “1st Configuration”, we modified the camber
7

K. Yoshida and Y. Makino
near leading edge. This refinement was performed by removing the load deficit near the
leading edge. We used a simple quasi-inverse method for this purpose. This inverse method
was formulated by two dimensional supersonic linear theory and applied to wing section
every 5% semispan location [3, 11]. The target load distribution was re-calculated to treat the
effect of body on warped surface using Middleton and Lundry’s method [12], which was an
extension of the Carlson’s method. Furthermore, the chordwise thickness distribution was
replaced with the distribution of NACA 66-series airfoil, which was one of laminar airfoils at
low speed, because it was found to have also better transition characteristics in supersonic
flows [6]. The re-designed configuration was called “2nd Configuration”.
(c) Application of natural laminar flow wing design
In designing the NLF wing, we originally
developed a complete three-dimensional inverse
method to achieve the optimum pressure
distribution [6]. It consists of both CFD analysis
and geometry modification procedure. The
governing equation of the modification procedure is
based on supersonic lifting surface theory. A panel
method was used to solve the equation [9].
Figure 7 shows a flowchart of the inverse design
procedure. The target pressure distribution on upper
surface was the optimum pressure distribution for
the NLF wing. The target pressure distribution on
lower surface was estimated by subtracting the
optimum load distribution by the Middleton and
1. Target upper Cp
C p
Transition analysis
by e N method
2. Optimum load
ΔC p
Warp Design
by Carlson Method
C p
3. Target Cp
4. Inverse Method (NS
+Lifting Surface Th.)
5. Designed Geometry
final
initial
Figure 7. NLF Wing Design Procedure
Lundry’s method from the target pressure distribution on upper surface. In addition, the “2nd
Configuration” was used as an initial configuration.
First of all, pressure distribution on the wing surface of the initial configuration was
estimated using a JAXA’s CFD solver with turbulent flow condition. Secondly, the pressure
difference between the estimated and the optimum pressure distributions was calculated.
Thirdly, the increment of geometry was estimated using the pressure difference and the
inverse method. Fourthly, the wing section geometries at 14 spanwise stations were modified
by adding the increment to the initial configuration. Finally, new wing geometry with
fuselage was re-defined by smoothing process using a three-dimensional geometry generation
software CATIA. These steps composed one iteration loop. Since we spent most of time for
the smoothing process in one iteration loop, it was required about one week for one iteration
including CFD analysis.
According to the design procedure, we conducted this iteration loop [3, 11]. In each
iteration loop, we experienced thicker modification at inboard wing and thinner modification
at outboard wing. Such modification was not corresponding to the prescribed spanwise
thickness ratio distribution. This meant that our NLF wing design method was not well-posed.
Exactly speaking, our target pressure distribution was not satisfied with both warp condition
and prescribed spanwise thickness ratio distribution. Although the target pressure distribution
should be improved to solve this problem mathematically, it is not easy to find other desirable
8

K. Yoshida and Y. Makino
pressure distribution to delay transition satisfying the conditions mentioned above.
Therefore, as an approximation, we adjusted the maximum thickness of the modified wing
geometry to the prescribed maximum thickness at the smoothing process by CATIA. The “3rd
Configuration” was designed after six iterations even though no complete convergence was
obtained. However, desirable transition characteristics were not obtained using the SALLY
code. The main reason was originated in the resolution of CFD analysis near the leading edge.
Therefore, in the re-design phase, the resolution was improved to realize the pressure gradient
almost same as that of the target one in accelerated region. Furthermore, we kept the
prescribed thickness ratio at outboard wing because of severe structural design requirements.
However, we did not adjust the maximum thickness of the modified geometry at inboard wing
to approach the convergence. Finally, we selected the modified wing configuration after ten
iterations as the “4th Configuration”. Consequently, the “4th Configuration” had the
improvement of the pressure gradient near leading edge as we already expected. The
transition analysis of the “4th Configuration” by the SALLY code indicated desirable
transition characteristics.
The wing section geometry of the “4th Configuration” was shown in Figure 8 and its
spanwise thickness ratio distribution was also shown in Figure 5. No adjustment of its
maximum thickness to the prescribed one was reflected in thicker distribution at inboard wing
region. And this led to a remarkable deviation from the exact supersonic area distribution due
to the Sears-Haack body. Naturally, it meant an
increase of wave drag due to volume. However, we
recognized the validation of the NLF wing concept
was more valuable than the validation of high L/D
in flight test.
The SALLY code is formulated in the
framework of incompressible stability theory.
Therefore, in the analysis of the supersonic NLF
wing, we can not predict transition location
quantitatively, including little database for
transition criteria on the so-called N value.
Furthermore, any codes based on compressible
stability theory are not available in Japan.
Therefore, at least, to solve the formulation
problem, we developed an original
compressible e N code, which was called
LSTAB code [13]. And we also tried to
develop a reliable transition database on the
critical N value for onset of transition in
supersonic flow.
Transition characteristics of the “4th
Configuration” were estimated using the
LSTAB code as shown in Figures 9. Here we
assumed N=14 as the criterion of transition,
Y/L
0.0
-0.1
8*z/l
8*(z/L)
0
-0.01
-0.02
-0.03
NEXST-1: NLF Wing
0.3 0.4 0.5 0.6 0.7 x/L
x/l
Figure 8. Designed NLF Wing Geometry
M=2.0, H=15km, N TR. =14
-0.2
α=0.0°
α=1.0°
α=1.5°
-0.3 α=2.0°
α=2.5°
α=3.0°
DP
-0.4
Estimated turbulent region
HF
TC influenced Cp(UPACS: by AS2-grid, the attachmentline
LBL(Kaups-Cebeci), contaminationLSTAB(Path B)
X/L
TBL condition),
Preston
-0.5
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0
Figure 9. Estimated transition locations at H=15km
9

K. Yoshida and Y. Makino
referring to the NASA’s test results at “Low Disturbance Supersonic Tunnel” [14]. As shown
in Figure 9, we predicted larger laminar region on upper surface at both design angle of attack
2 degrees and at slightly high angle of attack 3 degrees at flight test conditions, than that of a
usual SST wing. Therefore, the “4th Configuration” was selected as the final configuration of
the NEXST-1 airplane.
Furthermore, there is a possibility of transition due to attachment-line contamination in
flight test. It originates in turbulent boundary layer on fuselage surface. According to Poll’s
criterion [15], the transition due to attachment-line contamination was predicted at small
spanwise region of inner wing as shown in Figure 9. The validity of the prediction will be
judged by the measurement of transition
locations using the following techniques:
hot-film (HF), dynamic pressure transducer
(DP), thermo-couple (TC), and Preston tube.
2.3 Evaluation of Present Aerodynamic
Design Technique
We evaluated the effectiveness of
applying present aerodynamic design
technology developed for the NEXST-1
airplane. The principal results were
summarized in Table 1 and Figure 10 [16].
In the table, total drag was approximated by
the following expression:
2
C = C + C + K C − C + ∆C
D
Df
Dw
Mdes=2, CLdes=0.1 Concorde NEXST-1 Real 2nd SST
1. Pressure Drag Estimation by JAXA-CFD code
K 0.526 0.429
CL0 (CFD) 0.0171 0.0067
CDi’=K(CLdes-CL0) 2 0.003615 0.003734
CDw’=CDw+ΔCDi 0.004859 0.004063 0.003625 (dB, lTail↓)
CDp(=CDi’+CDw’) 0.008474 0.007797 (8.0%↓) 0.007359 (13.2%↓)
2. Friction Drag Estimation by JAXA-CFD code with correction
ReMAC(H=15 km) 174×10 6 22×10 6 202×10 6
Laminarization 0%(Turb.) 0%(Turb.) 60%(up.) 0%(Turb.) 30%(up.)
CDf 0.003393 0.005807 0.004990 0.003626 0.003336
(71.1%↑) (47.1%↑) (6.9%↑) (1.7%↓)
3. Lift-to-Drag Ratio Estimation
CD0(=CDf+CDw’) 0.008252 0.009870 0.009053 0.007251 0.006961
CD(CD0+CDi’) 0.011867 0.013604
(24%↑)
0.012787
(7.8%↑)
0.010985
(7.4%↓)
0.010695
(9.9%↓)
L/D 8.427 7.351
(12.8%↓)
7.820
(7.2%↓)
9.103
(8.0%↑)
9.350
(11.0%↑)
Table 1. Comparison of estimated L/D characteristics
(
Ldes L0
)
Di
≡ CD0
+ C′
Di
( 1)
2
+ C + ∆C
, C′
= K( C − C ) , C ≡ C′
+ ∆C
( 2)
where CD0
= CDf
Dw Di Di Ldes L0
Di Di Di
Here C Df , C Dw and C Di are friction drag, wave drag due to volume and lift-dependent drag
respectively.
In the application of the NLF wing
C L
Polar curve at M=2 and H=15km
design technique, we also assumed 60%
0.15
laminarization on upper surface for the
0.1
NEXST-1 airplane and 30% laminarization
C Ldes. =0.1
W/T(8.5% 風 試 模 型 )
for a real second generation SST
○:W/T(8.5%Model)
CFD( 実 験 機 )
0.05
▲:NEXST-1
configuration. The 60% laminarization was
CFD( 想 定 実 機 )
■:Real 2nd SST
◆:Concorde-like
CFD(コンコルド:
without
推 進 系
nacelle
無 し)
almost confirmed in our design study
--- : Concorde-like CFD(コンコルド: with 推 進 nacelle 系 有 り)
0
mentioned above. The 30% laminarization 0.006 0.008 0.01 0.012 0.014 0.016 0.018 0.02 0.022 0.024
was also predicted if we used a new
C D
optimum pressure distribution for the NLF
-0.05
wing design at high Reynolds number
condition. It was derived in our fundamental
Figure 10. Comparison of estimated polar curves
L/D=9.4
NLF effect
L/D=9.1
L/D=8.4
L/D=7.4
research activity. Furthermore, as mentioned above, since the tail cone length (l Tail ) and the
diameter of fuselage (d B ) of the NEXST-1 airplane were not similar scale to the real second
10

K. Yoshida and Y. Makino
generation SST, we used an original
configuration for evaluating the real second
generation SST.
Consequently, we can obtain that the
improvement of the L/D of a real large SST by
applying present aerodynamic design technique
is about 11% at cruise condition, compared with
the L/D of Concorde (see Table 1). Figure 10
shows estimated drag polar curves of some
configurations. We can see a remarkable
improvement on the real second generation SST.
Here “Concorde” in the Table 1 dose not mean
the exact Concorde configuration. It
ONERA-S2MA : IR Image:M=2, Re MAC =4.7×10 6 (P 0 =0.6 bar)
Flow
23.3% model
Hot-film sensors
IR camera
Transition line
Design Point:α=2°
Transition line
Off-Design Point:α=-1°
Figure 11. Transition test of NEXST-1 at ONERA
corresponds to a Concorde-like configuration without nacelle, which was designed by JAXA
according to the design philosophy described in other technical papers.
2.4 Wind Tunnel Tests
To complete the aerodynamic design of the NEXST-1 airplane, several wind tunnel tests
were conducted and the detailed results are summarized in reference [17]. In this section, the
principal results on confirming the NLF wing design concept are summarized below.
In general, it is not easy to conduct any transition measurement tests at usual blow-downtype
supersonic wind tunnels, because of the not-so-small freestream turbulence. However,
we estimated relatively lower freestream turbulence level due to continuous flow than that of
any blow-down-type supersonic tunnels. Therefore, we conducted the transition measurement
test at the supersonic closed-circuit-type tunnel of ONERA. It was called “S2MA” and had
the largest test section among several tunnels available for us. We used a wing-body
configuration test model with 23.3% scale of the NEXST-1 airplane (see Figure 11).
The infra-red (IR) image technique as well as multi-element hot-film sensors was utilized
in detecting transition location. The total pressure fluctuation measured by JAXA was about
0.29% using an unsteady pressure transducer placed on the model. Even though this value
was not so small, we clearly confirmed the rearward movement of the transition location at
the design point condition (angle of attack 2.0 degrees) as shown in Figure 11 [18]. However,
the amount of the transition movement was not the same as the predicted result by our
LSTAB code because of the freestream turbulence. Therefore, we are expecting that complete
validation will be conducted in the flight test.
3 AERODYNAMIC DESIGN OF NEXST-2 AIRPLANE
The jet-powered experimental airplane (NEXST-2 airplane) program is the second step of
developing advanced design technologies. In the aerodynamic design of the NEXST-2
airplane, the main target was placed on reducing both supersonic and subsonic drag. We
selected the YJ-69 by Teledyne as a candidate engine for the NEXST-2 airplane, because of
the limitation of cost and availability [19]. However, it had larger maximum diameter than
11

K. Yoshida and Y. Makino
that of fuselage. Therefore, we can not validate high L/D characteristics in flight tests because
of strong interference drag between the airframe and two nacelles.
In general, reducing such interference drag is the most important target in a practical
aerodynamic design. Therefore, we developed an original CFD-based optimum design system
for reducing the interference drag and the nacelle drag. It consists of both a JAXA’s Euler
solver with an overset grid method for calculating flow characteristics and an adjoint method
for analyzing sensitivity of objective function with respect to design parameters [20, 21]. It is
well known that the adjoint method is very effective comparing with the finite difference
sensitivity calculation using CFD.
Furthermore, we selected the decreased design Mach number of 1.7 because of reducing
supersonic drag and the limitation of the engine performance. The dimensions of the NEXST-
2 airplane were similar to the NEXST-1 airplane. In our flight test plan, the NEXST-2
airplane must be accelerated from subsonic to supersonic speed using own two jet engines.
Therefore, it is necessary to suppress the rapid increase of drag near the sonic, that is, M=1.0.
Finally, it is very important to design an optimum intake. JAXA conducted a lot of wind
tunnel tests on supersonic intakes and developed some useful design tools including a CFDbased
analysis code. Their research activities are summarized in references [22, 23, 24].
3.1 Design Concepts
The design concepts incorporated in the
aerodynamic design of the NEXST-2 airplane
are summarized in Figure 12. They are an arrow
planform with a little higher aspect ratio, a
warped wing with supersonic leading edge at
outer wing, an NLF wing at inner wing only, a
non-axisymmetrical area-ruled body and an
optimum nacelle. Two latter concepts were
achieved using our CFD-based optimum design
system. In particular, the non-axisymmetrical
area-ruled body concept was very effective to
reduce interference drag due to large nacelle
Design point : C L =0.1 @ M=1.7
Supersonic L.E.(Λ LE =42°)
Subsonic L.E.(Λ LE =66°)
Non-linear and
non-axisymmetrical
area-ruled body
Carlson’s warp
NLF wing (inner)
2.5th Configuration
Arrow planform
・AR=2.4
・S=10.12 m 2
under the wing, because the lower geometry of fuselage was dominant in controlling the
interference drag.
The remarkable difference of the aerodynamic design between the NEXST-1 and the
NEXST-2 airplane was to consider intake, diverter and internal flow of nacelle. The intake
and nozzle of the NEXST-2 airplane were designed using CFD analysis and experimental
study by JAXA’s propulsion design team. They found that an original intake shape of external
compression type and a convergence and divergence (CD) nozzle were very effective. The
diverter was also designed by both the propulsion and aerodynamic teams using the CATIA
system in the process of trial and error. Tail configurations and positions were changed to
satisfy flight controllability. In particular, the areas of horizontal tail and vertical tail wings
were set to 1.3 times of the areas of the NEXST-1 airplane.
Finally, we had to improve the precision of analyzing the flowfield around the NEXST-2
YJ-69
12 m
Optimum
nacelle
4.93 m
1.3×S NEXST-1
CD nozzle
Figure 12. Design concepts of NEXST-2 airplane
12

K. Yoshida and Y. Makino
airplane. The main part of the improvement was based on how to simulate the operation
condition of engine, that is, how to estimate inflow into the intake and outflow from the
nozzle as precisely as possible. In general, we often analyzed the flowfield around the
airplane using the so-called flow–through-nacelle condition as an approximation. To validate
this approximation was one of the targets of developing advanced design technologies.
1) Concepts related to the aerodynamic design of NEXST-1 airplane
In order to improve subsonic aerodynamics, a higher aspect ratio wing was desired after
designing the NEXST-1 airplane. Keeping the same kink position of both leading and trailing
edges for simplicity, we selected a modified arrow planform with the aspect ratio of 2.4 with
0.2 higher than that of the NEXST-1 airplane. However, the increment of the aspect ratio led
to reducing the swept angle of outer wing. Therefore, the outer wing had a supersonic leading
edge and every section had a pointed leading edge to reduce the profile drag.
A warped wing was designed according to the same design procedure as the NEXST-1
airplane without nacelle condition. The NLF wing design was applied at inner wing only
because the outer wing section had a pointed leading edge. The shpae of the target pressue
distribution was the same as that of the NEXST-1 airplane, but the pressre level was different.
In general, it is difficult to incorporate any influence of shock waves due to the nacelles into
the target pressure distribution. Therefore, the design C L for warp design was specified by
removing the increment due to compression lift generated by two nacelles from the total lift.
JAXA’s CFD calculatons estrimated that the increment of compression lift was about 0.06 at
the total lift of C L =0.1. Therefore, the NLF wing design was conducted at low lift condition,
that is, C L =0.04 without nacelles. In the design process, geometry modification was also
performed for the simple configuration without nacelles, but CFD analysis for estimating flow
characteristics was conducted for a complete configuration with nacelles.
2) New design concepts
An axisymmetrical area-ruled body for the NEXST-2 airplane was also designed by the
same design procedure as the NEXST-1 airplane under the the following constraints:
minimum diameter, minimum volume and maximum length. These constraints made it
difficult to apply the exact area-ruled body concept based on linear theory because of the
existence of two large engine nacelles. In addition, the design concept besed on linear theory
can not treat the influence of shock and expansinon waves due to intake and nacelle. Therfore,
the area-ruled body concept need to be expanded to include any nonlinear effects.
The key point of the non-axisymmetrical area-ruled body concept is to optimize the upper
and lower geometries of fuselage independently, using a certain mathematical fomulation of
two kinds of radial distributions as design variables. And total pressure drag was chosen as an
objective function. In addition, the overset grid method is also very useful in optimizing a
complicated wing-body configuration with such large nacelles. The optimum configuration
designed using our CFD-based optimization method was storngly influenced by the initial
configurtion. We designed the initial configuration using present aerodyanimc design
technique of the NEXST-1 airplane.
In the develpment phase of the method, we designed a non-axisymmetrical optimum area-
13

K. Yoshida and Y. Makino
ruled body for the “0-2nd Configuration” which
was defined in the conceptual design phase
described in the following section as a test case [20].
Figure 13 shows the designed fuselage
configuration, compared with the area-ruled body
based on linear theory of the “0-2nd Configuration”.
The influence of the nacelle under the wing leads to
the non-axisymmetrical feature on the fuselage
geometry. Therefore, this concept is very effective
to reduce the interference drag of such a
complicated configuration with relatively twin large
nacelles.
We also investigated an optimum nacelle
configuration using the optimum design system. Since the design space of the nacelle was
relatively small, that is the forward part and total length of the nacelle had to be fixed, we can
not obtain remarkable drag reduction effect for the optimum designed nacelle configuration.
However, comparing the initial nacelle and the optimum nacelle, we found some important
knowledge of reducing the drag. Furthermore, we investigated the effect on the nacelle
position, that is, streamwise station and spacing of two nacelles. We found the optimum
spacing of the nacelles [21].
Consequently, it was difficult to reduce the interference drag between the airframe and
nacelles drastically even though applying the CFD-based optimum design method because of
relatively large nacelle configuration. In genral, the principle to reduce the interference drag
was to decrease the cross sectional area of total configuration. One of the best ways was to
embed the nacelles into the wing under the limitaiton of structural design constraints.
3.2 Design Process and Results
Our aerodynamic design process consisted of two phases. The first phase was the
conceptual design phase. The main objective was to compromise aerodynamic performance
with some constraints required in developing a practical aircraft such as wing position,
aerodynamic tail-volume, intake geometry, front part of nacelle configuration, and embedded
nacelle configuration. The next design phase was aerodynamic optimum design phase. We
focused on the non-axisymmetrical area-ruled body design, including the optimum nacelle
design. In addition, the non-axisymmetrical area-ruled body concept has large possibility to
be extended for reducing sonic boom. This is investigated in our fundamental research
activity [25]. Design results at each process are described in the following sections.
1) Conceptual design phase
We designed a fundamental configuration with two large nacelles as a first step. This
configuration had a simple straight body with a Karman ogive cone and a warped wing by
Carlson’s method. The planform was selected considering drag characteristics using Carlson’s
method according to the same selection rule as the NEXST-1 airplane. The remarkable
difference was to adopt supersonic leading edge at outer wing as stated above. This
Fuselage radius
Linear area-ruled Body
Non-linear optimized
area-ruled Body
(r/L)
0.03
0.02
0.01
0.00
0.0
0.2
Initial
Upper
Side
Lower
0.4
x
0.6
Axisymmetric
*0-2nd Configuration
Non-axisymmetric
ΔC Dp
=-0.0006@M=1.7
Figure 13. Comparison of optimum arearuled
body design
0.8
1.0 x/L
14

K. Yoshida and Y. Makino
configuration was called “0-1st Configuration” of
the NEXST-2 airplane.
Figure 14 shows a wind tunnel test model for
the “0-1st Configuration”. It had a flow-plug just
after two nacelles and several total pressure
measurement tubes placed near exit station of the
nacelle. We were able to control the inflow
entering into the intake by changing the plug
position. We obtained useful experimental data
using this model, as mentioned later.
In the conceptual design phase, the “0-1st
Configuration” was the start point. As the next
step, we designed from the “0-2nd Configuration” to “0-8th Configuration” considering an
area-ruled body concept based on linear theory, some improvements of nacelle configuration
and its position, optimization of intake geometry, and aerodynamic tail-volume. This design
process was not straight-forward but a process of trial and error approach in reducing
aerodynamic drag within several constraints required in practice.
2) Aerodynamic optimum design phase
(a) Baseline configuration design
The “0-8th Configuration” was only designed from the aerodynamic viewpoint. To
develop a practical experimental airplane, it was necessary to compromise the following
fields: airframe aerodynamics, structural constraints, flight dynamics, intake aerodynamics,
and propulsion system performance. As the first step, we designed the “1st Configuration”
using present CFD-based aerodynamic design method, maintaining several practical design
constraints required by industries.
In the aerodynamic design of the “1st Configuration”, the NLF wing design at inner wing
was conducted using the CFD-based inverse design method and a modified target pressure
distribution. Figure 15 shows the estimated laminar region using our compressible e N code. In
this estimation, we selected the transition criterion
of N=10 as a reference. This lower criterion than
the N=14 for the NEXST-1 airplane was assumed
by taking account of the influence of embedded
nacelle on the upper surface such as acoustic
disturbance.
Two nacelle configurations and their positions
were optimized under the design condition of
intake and diverter, using the CFD-based
optimum design method [21]. In addition, a nonaxisymmetrical
area-ruled body was designed
using the same method [20]. However, we did not
obtain the remarkable drag reduction effect of the
concept. We considered that present small
Flow-plug
Figure 14. W/T model for “0-1st” Configuration
of NEXST-2 airplane
y/l
0.5
0.4
0.3
0.2
0.1
0
-0.1
NEXST-2: Estimated Transition Position
M=1.7 , α=0.5°, H=15km
Wb0803aGeometry contours
Estimated laminar region
N=2 4 6 8 10
0 0.2 0.4 0.6 0.8 1 x/
Figure 15. Transition Analysis on NEXST-2
(1st Configuration)
15

K. Yoshida and Y. Makino
reduction was dominated in selecting the initial configuration, that is, the “0-8th
Configuration”.
(b) Improved configuration design
After investigating the “1st Configuration” from several technical viewpoints by the
industries, new requirements for its flight tests were specified as follows: 0.5m extension of
fuselage, more fuel capacity, increase of thickness ratio from t/c=3% to 5% at outer wing,
about 2 inches larger diameter of the fuselage, about 0.1m size-up of the nacelle diameter, and
keeping the clearance of 0.015m between the nacelle and lower surface of the wing at front of
the diverter. Furthermore, the pressure drag of the NEXST-2 airplane near the sonic speed had
to be drastically reduced, compared with the drag characteristics of the “1st Configuration”.
Therefore, reducing the drag at the whole Mach number range was the most important target
in the improved design phase.
In the aerodynamic design of an improved configuration, first of all, previous NLF wing
was adopted without any modification. And the optimum design of nacelle shape and position
was performed according to the usual approach. The optimization of nacelle position was
carefully conducted, and an optimum solution was obtained. On the other hand, although an
optimum nacelle shape was certainly found, it was limited by lower bound of the structural
constraints. However, the comparison of the optimum shape with the initial shape provided us
important aerodynamic insight to modify the shape. Therefore, as for the nacelle shape, we
modified the optimum shape using the CATIA system according to the aerodynamic insight
under the relaxed structural constraints. Figure 16 shows the principal result for optimizing
the nacelle shape.
As the first attempt of improving the
fuselage of the “1st Configuration”, the
non-axisymmetrical fuselage design
concept was applied to design the rear part
of the configuration, taking account of the
influence of the horizontal tail. In this
design process, the initial configuration
was the “1st Configuration”. Consequently,
we defined the combination of the
designed fuselage, the CATIA-based
optimum nacelle and the NLF wing. This
combination was the “2nd Configuration”.
Furthermore, to reduce interference
drag near the sonic speed, a new optimum
(a) Initial (b) Optimized (c) CATIA Final
Non-axisymmetrical Area-ruled ruled body Design
non-axisymmetrical area-ruled body was designed again using the different initial
configuration. We considered the initial configuration with similar area distribution to that of
the “0-7th Configuration”, because we found the “0-7th Configuration” had a potential to
reduce interference drag near low supersonic speed, that is, M=1.2. Figure 16 also shows
design variables on the optimum design procedure and the designed non-axisymmetrical arearuled
body. The designed configuration was called the “2.5th Configuration”. Finally,
R_upper
Initial
R_lower
Optimum Design of Nacelle shape
R_side_upper
R_side_lower
Design variables
Side View
Top View
Initial
Optimized
Wing Upper
Wing Lower
Figure 16. CFD-based optimum design of NEXST-2
16

K. Yoshida and Y. Makino
minimum pressure drag characteristics were
estimated over the whole Mach number range as
shown in Figure 17. We made best use of the nonaxisymmetrical
area-ruled body concept and
obtained remarkable improvement in aerodynamic
performance. Consequently, the “2.5th
Configuration” was selected as the final
aerodynamic configuration for the NEXST-2
airplane.
2.0th Configuration
2.5th Configuration
3.3 Wind Tunnel Tests
Figure 17. Minimum pressure drag of NEXST-2
To understand the strong airframe/nacelle
interference completely, and to develop the reliable and effective database for supersonic
intake, several wind tunnel tests were conducted [22-24, 26-27]. In this section, principal
results on fundamental research activity except intake tests are summarized below.
1) CFD validation tests of the airframe/nacelle
interference configuration model
The test model of the “0-8th Configuration” with
flow-through-nacelle was used to conduct the CFD
validation for a complicated configuration with
nacelles. Figure 18 shows the force test model with a
modified intake shape and a CFD grid. The scale of
the test model was 8.3% of the NEXST-2 airplane.
Both supersonic and subsonic tests were conducted at
JAXA [26]. In the test at flow-through-nacelle
condition, the drag measured by a force balance
should be corrected using the relation of momentum
balance of internal flow of the nacelle.
Figure 19 shows measured and corrected drag
polar curves, comparing them with CFD(NS)
computation results. An open circle and triangle
symbols indicate test results of drag characteristics
without and with the internal flow correction
respectively. Our CFD computation results on both
conditions without and with the internal flow
correction are fairly smaller than those of the test
results. It strongly depends on the turbulence model in
our CFD code. If we assumed the drag characteristics
with an off-set value 0.0049 of the minimum drag, we
found very high correlation between those assumed
5.06 million points:
supersonic flow case
Figure 18. W/T model & CFD grid of
“0-8th configuration”
0.01 0.02 0.03 0.04 0.05 C D
drag polar curves and test results. Therefore, the pressure drag characteristics estimated by the
CFD code were well validated.
CL
0.4
0.3
0.2
0.1
0
-0.1
NEXST-2-08 : Wing+Body+Nacelle
M=1.7
No Int. Correc.
With Int. Correc.
BWT(20099) NS Off-set(0.0049)
Figure 19. Comparison of test and CFD
results
17

K. Yoshida and Y. Makino
2) Flow-plug tests for the nacelle flow effect on drag characteristics
In order to collect the complete drag data of the NEXST-2 airplane in the system design
phase, we had to investigate the mass flow effect of
the nacelle on the airframe drag characteristics.
Therefore, we produced a large test model with two
flow control plugs behind each nacelle shown in
Figure 14. This model was a 17.0% scaled model. A
flow plug test was conducted in the 2m x 2m
transonic wind tunnel of JAXA [27].
Figure 20 shows drag characteristics with respect
to mass flow ratio (A0/Ai) of nacelle flow at Mach
1.4. Here A0 and Ai mean the cross-sectional area of
actual flow stream tube at forward infinity and the
capture area at the front of the intake. The figure
indicates an increase of the nacelle drag as the mass
flow ratio decreases. The total drag strongly depends
on the nacelle drag. Figure 21 shows pressure
distributions measured by the pressure sensitive paint
(PSP) technique [28] at the typical two cases of small and
large values of A0/Ai. As shown in the figure, the strong
shock wave was observed in front of the intake at small
mass-flow ratio condition. This flow pattern corresponds
to the unstart condition. On the other hand, at the
condition of large mass flow ratio, there was no
remarkable shock wave in front of the nacelle.
4. CONCLUDING REMARKS
Nacelle flow effect : M=1.4, α=0°, β=0°
CDFc
0.03
C D 0.025
0.02
Wing-Body
0.015
0.01
Nacelle(Left)
0.005
0
-0.005
0 0.2 0.4 0.6 0.8 1
Ao/Ai(Left)
A 0 /A i (L
Figure 20. Interference test of NEXST-2 :
JAXA developed some original advanced design concepts and procedures in the
unmanned and scaled supersonic experimental airplane program. The supersonic NLF wing
design concept and the CFD-based inverse design procedure were developed for the first
airplane. The non-axisymmetrical area-ruled body concept and the CFD-based optimum
design procedure were developed for the second airplane. The NLF wing concept was
validated in the wind tunnel test qualitatively, but not quantitatively, because of the existence
of freestream turbulence in any supersonic wind tunnels. The flight test is expected to validate
it both qualitatively and quantitatively.
In reducing strong interference drag between the airframe and two large nacelles of the
second airplane, the effect of the non-axisymmetrical area-ruled body concept was confirmed
numerically. However, the concept has not been validated experimentally, because it is not
easy to simulate the complete flowfield around a complicated configuration with engine
operation condition in wind tunnel tests. The flight test is valuable in validating the design
concept. We expect to continue to develop the advanced aerodynamic design technology after
the successful flight test of the NEXST-1 airplane.
0.045
0.04
0.035
Total
Plug test (“0-2nd Configuration”)
Nacelle flow effect : M=1.4, α=0°, β=0°
Ao/Ai=0.33 Ao/Ai=0.86
Figure 21. Interference test of NEXST-2:
PSP test (“0-2nd Configuration”)
18

K. Yoshida and Y. Makino
ACKNOWLEDGEMENT
The authors would like to express special thanks to the Mitsubishi Heavy Industries,
Kawasaki Heavy Industries, Fuji Heavy Industries and Tohoku University for their
cooperative efforts to the design of experimental airplanes. In addition, a lot of research
activities related to those designs were supported by some researchers of JAXA. The authors
also would like to thank them.
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