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NASA-TM -88444
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NASA TECHNICAL MEMORANDUM NASA TM-88444
INITIAL ADAPTATION TESTING OF THE BIDIMENSIONALLY SELFADAPTING WALL OF TIm FRENCH T2 WIND TUNNEL, AROUND A THREEDIiMENSIONAL OBJECT:'
J.P. Archambaud, J.B. Dor, A. Mignosi and L. Lamarche
Translation of: "Premiers essais d'adaptation des parois auto-adaptables bidimensionnelles de la soufflerie T2 autour d'obstacles tridimensionnels." Rapport technique OA 33/3075 (DERAT 11/5015 DN), O.N.E.R.A., Centre d' Etudes et de Recherches de Toulouse, France, September 1985, pp. 1-49.
JUL. 1 4- 19BG
LANGLEY RESEI~RCH CENTErl LlBRAH't. NAS,~
Hi" '::>TfJN. VJRcrNIA
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON D.C. 20546 JUNE 1986
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4. Till ....... "'btUI. INITIAL ADAPTATION TESTING OF TIlE BIDIMENSIONALLY SELF-ADAPTING WALL OF THE FRENCH . T2 WIND TUNNEL~ AROUND A THREE-DIMENSIONAL OBJECT
J.P. Archambaud, J.B. Dor, A. Mignosi and L. Lamarche
s, n.p •• t 0.'0 JUNE 1986
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Translation
U. ~1_mIQl" 1409911 Translation of:"Premiers essais d'adaptation des parois auto-adaptables
d t 1:. 1 'd" , 1" bidimensionnelles de la soufflerie T2 autour ou,stac es trl lmenSlonne s. Rapport technique OA 33/3075 (DE RAT 11/5015 DN), O.N.E.R.A., Centre d' Etudes et de Recherches de Toulouse, France, September 1985, pp. 1-49.
~---------------~------------.------~----------------.------------------~~----,~
The test series were carried out at ONERA/CERT at the T2 wind tunnel in September 1984. The obj ective of this series was to minimize wall interference through a bidimensional adaptation around the 'models, inducing tridimensional flows. For this, we used three different models by measuring either the pressures or the forces and moment of pitch (balance). The adaptation was derived from a correction computation in the compressible axisymmetric tridimensional.
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TM88444 TOULOUSE ENGINEERING AND RESEARCH CENTER
2 Avenue Edouard Belin--31055 TOULOUSE CEDEX
DEPARTMENT OF ENGINEERING AND RESEARCH
IN AERODYNllJYIICS
~_ngine~ring R(~port OA 33/3075 AND (DERAT 11/5015 DN)- Sept 1985
First tests of adaptation of bidimensional a~aptive walls in
the T2 wind tunnel around tridimensional obstacles.
by J.P. ARCHAHBAUD
/s/
D.E.R.A.T. Director
R.IvlICHEL
/s/
.J. 13. DOR
/s/
--1-
i\. LvII GNOS I L. LAMARCHE
/s/ /s/
AUTHon'S ABSTRACT
~t'hese 1:ests carried out in September 1984 consti tute the
first approach to adaptation of tridimensional flow. The
deformation of the walls remains bidimensional (as in the
preceding test series). 13ut the computation of these deform
ations is the result of a method of estimation of wall
interferences in the tridimensional.
Three models have been tested:
- C 5 rotation body
- model of the F 4 transport aircraft
- model of a delta wing canard aircraft
/3/ *
The adaptation around the C 5 body had the objective of
comparison with the results of a true tridimensional adaptation
(TU-Berlin) and tests in a large wind tunnel (NASA Arnes).
*Numbers in ~~~grn indicate foreign pagination.
_. 2--
page /4/ not translated /5/
PT
RC
TT
2D
3D
0<.
v.g.
h.g.
NOTATIONS
test section airflow width
model length
coefficient of length
coefficient of pitch
coefficient of drag
height of test section airflow
test. infinite tlach number
total pressure
Reynolds number of the flow (linked to C)
total temperature
Cartesian coordinates
orthonormalized reference
abbreviation for bidimensional
abbreviation for tridimensional
angle of incidence announced before the test
corrected angle of incidence (by balance measurements
or photographic measurements)
vertical generatrix
horizontal generatrix
-3-
NASA Ames test
/6/
1. IN'l'RODUC'l'tON
The test series presented below were carried out at ONERA/
CERT at the T2 wind tunnel in September 1984. It was carried out
in close collaboration with Prof. U. GANZER and his associates
MM. Y. IGETA and J.ZIEMANN at the University of Berlin.
The objective of this series was to minimize wall inter
ference through a bidimensional adaptation around the models,
inducing tridimensional flows.
For this, we used three different models by measuring
either the pressures or the forces and moment of pitch (balance).
The adaptation was derived from a correction computation in the
compressible axisymmetric tridimensional.
2. GENERALIZATIONS
2.1. T2 Wind Tunnel
The T2 wind tunnel is a closed loop installation with induct
ion operation wi·th 1 to 2 minute gusts (PL.I) /REF. I and 2/.
Its test section airflovv (H :: 370mm, B :: 390mm, at the inlet)
is equipped with flexible upper and lower walls (PL.I), @ach
activated by 16 jacks. Each deformable wall is equipped with 58 central
pressure ,Ports (¢ :=: O.4mm) and with some lateral ports. In addition,
the left vertical wall also has pressure ports arranged in three
horizontal and three vertical lines. Only the central ports will
be used during the adaptation; the others make possible verification
of satisfact.ory ooherence of the veloci ty field only iJ:). some cases.
2.2. ~odels--balance--sting model support
The three models used are;
- axisyrrrnetric C5 body c:~ 166. 2 SHun
-. 1"'4 aircraft C .- 119. 9mm
_. delta wing canard C .- 145mm
'1'he C5 body is an axirnetrique model formed of an assembly of
geometrically simple elements (PL. 2 and 3). The body has been
tested in numerous installations which makes possible the comparison
of interesting results; to wit: 3D adaptive walls TU Berlin--16 ft
AEDC--ll ft NASA funes, 6 ft 82 Modane ONERA.
-4--
The model called the F4 is one of an Airbus type aircraft
and has suparcritical wings (PL.4). The third model represents /7/
a supersonic ~elta winged aircraft equipped with two small
C:~~ard-·type _~:tilerons at the front.
A balance of very small overall dimension (¢ = 8mm) furnishes
the axial and norhlal loads and the pitch moment working on the
model. Knowledge of the loads and of the pitch moment also
makes~ssible definition of the actual angle of incidence. This
balance is adaptable to two aircraft models. The electrical voltages
associated with the different components are obtained in real
time and recorded by the acquisition system.
'.I~he various models are held by a sting support (see PL. 2) the
diameter of which creats a tiny perturbation on a length of about
a chord. Further downstream, this sting is connected to a rather
voluminous gimbal joint centered approximately 2 chords behind the
profile, towards the downstream end of the adaptive wallsi the
overall bulk of this obstacle will therefore be compensated by
the adaptation of the airflow section, thus avoiding a very strong
corresponding perturbation and its upstream extension very far
towards the model. The part at the rear of this gimbal is fixed to
a rod which lies across the second throat and pivots arownd an axis
of rot.ation l.5mm downstream of the model.
2.3 Adaptation of the upper and lower wall
The physical displacement of the adaptive upper and lower
walls is bidimensional (REF. 3/. But the deformations computed
by the program for adaptation /REF. 4/ take into account the
tridimensional dharacter of the flow. With this end in view, the
model is designed for the distributions of 3D sources and of horseshoe
vortices (of infinitely small dimensions) located on the airflow
axis. The pressure measurements on the upper and lower walls give
access to t~ in~nsities of these singularities. The interference
of the walls is thus estimated by means of external imaging at
the level of the airflow axis. It is next cancelled out by an
appIDpri~e new formation to the walls. Actually, two linear
operators, uniquely dependent on the test section geometry, make it
possible to go directly from parietal ~ffisure measurements to the
adaptive forms. The compressibility is taken into account by
-5,- ..
the preSenCE! of the factor 11 == 1i--- Bo2. 'rhe method thus des
cribed consists of a single iteration.
The rel~asing thus created around each of the three models is
not very important. due to the small overall dimension of these
obstacles in the T2 test section. On the other hand the sting
model support with its relatively voluminous gimbal (PL.2)
causes a strong divergence of the adaptive walls which is
frequently insufficient (lower wall in extreme position--Plate 7).
However, in all the cases presented the perturbations caused by
the model and the sting seem well decoupled in the wall area: either
because these perturbations are initially weak and little extended
between non-adapted walis (M ~ 0.7), or because the adaptation o . diminishes and localizes the sting perturbation in the more
rigorous configurations.
In the following part of this report we will term "non
adapted walls" a simple divergent configuration designed to
compensate the convergence due to the boundary layers developing
on the four walls of the test section airflow (vertical walls
considered as plane plates /REF. 3/). The expression "adapted
walls" will designate the wall forms produced by the adaptation
computation in a stage described below.
2.4 Test configuration~
The test configurations are given in Tables 1, 2 and 3, and
are visualized in the following synoptic diagram:
-6-
/8/
Mo 1 -
IJ. C5 I
~ L
0 Fit 0.9 [
a ovi'on .t. cona,rd l u~
0.8
~l mest.;we Z pression l
CI
0.7
.. l us
610 il iil iI il I!I
:1 meSlJre 3, forces (balance)
0.6 C<
-2 o 2 4 6 8
1. canard aircraft
2. pressure measurement
3. measurement of forces (balance)
When the balance is not employed, the angle of incidence ~
examined in the plots is that measured before the run. Indeed
the rigid assembly does not prevent a slight variation in this angle
during the run . We have been supplied by Mr GANZER with this
incidence correction measured by means of photos; it is plotted
in Figure 32 and can serve to correct the corresponding plots in
this type test. On the other hand, when the balance is used, this
angle correction, systematically computed, is reflected in all the
corresponding results in this type test. Verification with a
cathetometer during run also makes it possible to check out
the satisfactory precision of this correction on several tests
carried out with the F4 aircraft model.
The test Mach number M is determined from the distributions o
of Mach number measured on the upper and lower walls. It involves
an average
slight for
(see PL. 5
in space. Local deviations with respect to Mo are
Mo ~ 0.85; these deviations then overlap D for M ;> 0.85; o 0
and 6).
For each test configuration (except for the C5 body, Mo = 0.70,
~ = 00) two runs are necessary. The first is carried out
.. --7-
o between non-adaptive walls, and makes use of the computational
base for the adaptive forms which will be used in the second run : /9/
this second test allows pressure and aerodynamic parameter measure
ments while minimizing the int.eractiOl1 of the walls.
All the tests were carried out at ambient temperatures with
a pressure generatrix between 1. 6 band 2.2 bar.
3 - C 5 BODY
The model (PL.3) provided by TU Berlin for the more reduced
test sections, presents a slight overall dimension in the T2
wind t:unnel (S . /s . ~ 3.10/00). It has
max cross sectlon test sectlon been tested at zero incidence, and the measurements carried out
are pressure measurments using 20 pressure ports.
The following table gives some accuracies in the cases covered:
T2 upp
and 10
wall
Compar
with
er
wer
ison
0.6
adapted
3.1. ~'Jalls
0.7
non-adapted
TU Ber-lin NASA Ames
-
0.843 0.915 0.93 0.95
adapted adapted adap·ted adapted
I TU Ber- NASA lin Ames NASA .. Ames
In plates 5 and 6 one confirms the faint signature of the
model on the flexible walls, as long as M < 0.85. Above this, a this influence is likely to be more and more significant.
In all the cases accomodated one notes the inadequacy of
0.97
adapted
NASA Ames
the unblocking of the test section foreseen by the correction
calculation for the site of the gimbal. This phenomenon is strongly
accentuated when M increases aboVe 0.85. However, all the tests a
show that this perturbation does not increase enough upstream to
interfere with those which are ascribable to the model.
The lack of parallelism of the walls appears in the form of
a modulation around the M value upstream of the model. The ampli-o
tude of this modulation grows with Mo. We are dealing here with
slight bulging of the wall between the jacks which seem to alternately
succeed one another in the direction of the test section and then
toward the exterior. 'l'hey are the results of a complex combination
-8-
-
of several parameters: the pressure differences between the interior
of the test section and the exterior container, upstream tail-in,
flexure of the sheet metal, and the relative position of "fixed"
points (rotation possible) at the site of the jacks.
Plate 7 shows the wall shapes. The central shape, simply
divergent, has been called the non-adapted shape. (see sec. 2.3.).
The other evolutions proceed from the adaptation computation .. One
sees a trend at the opening even before the first jacks, upstream
of the test section. Unblocking is very weak at the site of the
model. On the other hand, downstream the walls are s·trongly di
vergent but insufficientl~ as mentioned above. A certain dis
symmetry between the upper and lower shapes can be determined at
the sites of the two upstream rear jacks; indeed on the lower wall
the rearmost jack is at the extreme position in all cases and
also the next to last jack for M > 0.95. o Note : Plates 8 and 9 present the distributions of pressure
measured to the right of the model on three walls: upper, lower
and the left side.
In Plate 8, around the axisymmetric C5 body, the various
curves seem to be in good agreement except around x =- +IOOmm
(towards the base of the model) where a minimum Hach number is .. recorded more pronounced on the lower wall.
On Plate 9 (canard aircraft- ~= 80 --non-adapted walls) there
can clearly be perceived the continued diminution of the pressure
perturbation when it
test sec·tion airflow
evolution i 4'" .:;, coherent
passes through
by coming away
with the test
the demi-perimeter of the
from the upper wall. This
configuration.Moreover, the
symmetry with respect to the median vertical plane is'verified
correctly by the favorable cross-checking of the lines of
pressure ports situated on the edges of the flexible walls. However,
the pressure distributions on the downstream portion of the lower
wall x = 0 presents the same defect as in Plate 8, i.e. a trough
preceded here by an equally abnormal spike; these anomalies
recurred in a certain number of tests and seem attributable to
an imperfection in the sheet-metal.
It is interesting to note that these pressure distributions
recorded on the plane and rigid lateral wall flow much better
than those measured on the flexible walls.
'-9-
3.2. !>-1odel
Oil visualization carried out (PL.3) on the model at M = 0.6 o
shows that th~ laminar flow at the stagnation point (dark zone)
transitions before the central bulge (distinct cones of untimely
release).
Plates 10 and 11 show the distributions of Bach number on
the C 5 body in the different cases studied, (8xcept for
o • 7 <. f-1 < 0 • 84 ) • a
For M < 0.85 (PL.IO), the general shape of the run is prea served, whereas the maximum value around x/c = 50% perceptibly
increases with M . Plate 8 (M > 0.84) indicates a strong widening a a . of the median supercritical zone as M increases, on account of a the shock recoil~ in parallel, at the base of this shock, a
separation causes an overloading of the velocity minimum between
60% and 80% of chord.
Plates 12 and 13 make possible a comparison of the results
obtained wi t.h the same model at Berlin (TU Berlin- 2D and 3D
adaptation,/REF. 5/) and at ONERA/CERT (~2, 2D adaptation based
on a 3D correction). On the whole the cross-checks are good;
at M = 0.7 (PL.12) the results of the 2D adaptations are very o coherent, whereas the 3D adaptation (~O'eerlin) appears lo un-
block the more slightly in certain zones. Around M = 0.84 (PL. 13) a the Mach number distribution produced by TU Berlin (M = 0.84) o is well inserted between the two readings taken at '1'2 (M = a 0.832 and 0.843.
A comparison is also made between t:he results at NASA Arnes
/REF.6/ on plates 14,15,16 and 17 respE~ctively for ~101'= 0.7,
0.84, 0.95, and 0.97. The model used at Ames involves two rows
of ports drilled on two generatrices located in the perpendicular
plane. It is approximately 6.4 times longer and the transition
is released around about the stagnation point by ballottines.
The results are in good agreement; at H = .097, downstn-'!am of a
the central bulge the Reynolds effect seems to diminish the
intensity of the shock/boundary layer interaction in the case
of NASA Ames compared to T2.
3.3. Comparison of experiment and computation
Two cross-check tests 01 measurements, carried out at T2 by
-10-
/11/
~ , '.
computations were done, one on the model, the other on the walls.
The model has been schematicized by an aggregate of panels
of uniformly 'loaded sources i the velocity is calculated for
compressibility using the Goethert' s rule of similitude, valid for
subsonic circulating flow. A generatrix comprises 90 panels and
a transverse section 32. For the infinite Mach numbers considered
(M = -0.6- 0.7- 0.84), good experiment/computation agreement o
is established for the whole of the model (PL.18), mainly at
the extreme segments. The central spike is nicely reproduced by
the computation for M = 0.6 and 0.7; however a small deviation o
appears for M = 0.84 (sonic peak) and one could surmise a o viscosity effect and an imprecision in the similitude rule.
The model mount assembly (sting, gimbal, rod) has been
schematicized by a series of contiguous segments of linea~ized
doublets (constant intensity on each segment). The compressible
computation was also done to include Goethert' slaw. The images of
the doublet segments in relationship to four sides of the test
section schematicize the wall interference. Plate 19 shows the
distributions measured and calculated for the Mach number on the
upper and lower walls in two non-adaptive cases (r·l = O. 7 and 0.84); o one can confirm that this schematization predicts quite well the
perturbations of the sting and its gimbal.
4- 1"4 AIHCRA1"T
The F 4 aircraft model (PL.3) presents a very small overall
dimension in the test section airflow of the T2 wind tunnel
(S/SV = 2.5°/ ) and all the tests were carried out between non-00
adaptive walls. Plate 20 shows the faint signature of the model )
on the upper and lower walls in one of the cases studied which
was ·the configuration with the greater lift.
For this model we carried out a sweep in incidence at M -a 0.7. The aerodynamic coefficients CL,CD,CM are presented in
Plate 21.
5- CANARD AIRCRAFT
5.1 Tests with balance
These tests comprised an incidence sweep ( 0 ~ do- < 0.844)
at Mo = 0.70 and two sweeps in Mach number (0.7 < Mo < 0.844)
at 0<. =2 0 and 30 (selected).
Figure 22 shows the evolution of the lift coefficient CL
-11-
/12/
as a function of the incidence angle ~ at Mo= 0.7. There is
established a noticeable diminution of CL caused by the adaptation
of the flexible walls, a deviation growing '.vi th d... • Plate 23 regroups the curves CD (CL ) and CM (CL ) at Mo= 0.7
dCL
Plate 24 illustrates the increase in slope -- as I1 grows. d 0
On Plate 25 are plotted the values of CL as a function of 110
_I 0 0 ' for the initial incidences o~ ~ = 2 and 3 • The incidence
correction qi ven by the balance increases with 1110 (see table).
Reduced to a fixed incidence (incidence corrected at Mo= 0.7), the
evolutions of CL are quasi-rectilinear and very slightly increas
ing.
5.2. ~!!~sts with pressure measurements
The pressure ports number 10, of which 2 are located under
the cockpit. The 8 others form a line on the side of the fuselage,
above the plane of the wings.
This sE~ries of tests comprised a sweep in incidence
O~o<. < SO at M = 0.7 as well as a sweep in Mach 0.7 < M 0 0
< 0.876 at 0( =3 0•
In Plate 26 showing the distributions of Mach number on
the deformable walls, one clearly sees the growing influence of the
model on the upper wall when tIle incidence grows. The signature
on the wall is never significant. We note that the incidence
setting is made by rotation at the site of the gimbal, and
therefore the aircraft is located above the axis of the airflow
for (;l > 0 ..
One can note on the upper wall that the bulged shape of the
distribution of Nach number before adaptation dips in the middle
when this adaptation takes place. lVe also stress the perturbation
constant due to the gimbal but its interaction is more and more
marked with that of the model in the non-adaptive case.
On the model (PL.27), the velocities everywhere increase
with incidence, but in a more distinct manner between X/C =
50% and 75% at the wide part of the wings, the elements with the
most lift.
/13/
Up to 0(:= 60, the velocity curves on the model between non-
-12-
·"
adapfive and adaptive walls do not differ from each other. On the
other hand, for ~ =80 (PL.28) the non-adaptive test presents an
overspeed of ,the system on the order of A II[ = 0.01 in comparison
with the adapted case. It seems, from the velocity on the walls
(PL.25), that the non-adaptive case corresponds roughly to a real ,
infinite Mach number (linked to the proximity of the model)
slightly more elevated; this would explain the general displace
ment of 1,1ach number on the model i the effect of the incidence of
non-adapt ion would be of a lower order.
]\t a fixed incidence ( 0<. == 30), between non-adaptive walls,
one notices (PL.29) the phenomenon observed earlier of increase
and of the interaction of the perturbations of the model and
the sting/gimbal assembly when M increases. For high values of a
M , there results a longitudinal velocity gradient. The adaptation o of the walls creates unblocking of the test section in its
downstream section which has a higher level of velocity.
The longitudinal gradient has also disappeared and the
perturbations due to the profile and to the gimbal seem separated.
Plate 31 shows a regular staging of the Hach number distributions
on the profile, with a strong increase in level next to the
canard wings at the front (passage to supersonic) and on the •
cockpit.
CONCLUSION
This series of tests is the first step towards a minimizing
of the wall interferences in the tridimensional. The bidimensional
adaptation of the upper and lower v'lalls alone constitutes a pri~ri
the most rough approximation of the process; this shortcoming is,
however, reduced by the fact that the profiles studied' are of small
overall dimensions. On the other hand, the method is grounded in
a tridimensional computat.ion and a cancellation of the wall
interference on the test section airflow axis; this method,
applied to Cl "bidimensional" t.est section similar to that used
at T2, leads to a small residual interference according to these
authors /REF.4/.
--13-
/1/ HICHEL R.
QUEHARD C.
~lIGNOSI A.
/2/ GOBER~r J. L.
MIGNOSI A.
/14/
REFERENCES
The induction driven tunnel T2 of ONERA/
CERT : flow qualities, testing techniques
and examples of results.
-Journal of Aircraft, VOL. 16. No.3 (1979)
Studies on the cryogenic induction driven
wi nd tunnel '1'2.
-ETH cryogenic Technology Revie\,l Heeting
NLR Amsterdam (1982).
/3/ CHRVALLIER J.P. Adaptive walls at the T2 wind tunnel:
/4/
/5/
/6/
MIGNQSI A. principle, construction and some examples ARCHAMBAUD J.p.
SERAUDI A. LAMARCHE L.
WEDE~1EYER E.
GANZER U.
IGETA Y.
ZIE~1ANN J.
VAUCHERE'r X.
BAZIN M.
ARr-1AND C.
of bidimensional results.
-La Recherche Aerospatial No. 1983-4
Minimization of wall interference for
three-dimensional models with two
dimensional wall adaptation.
-V.K.I. Technical Note 149 (March 1984).
Design and operation of TU Berlin wind
tunnel with adaptive walls
-ICAS paper 84-2.1.1., sept. 1984.
Comparison of bi- and tridimensio~al
transsonic tests carried out in various
wind tunnels. -AGARD CP No. ]87 (]975)
-,14-
/15/ I
1- • Corps C 5 ( pression)
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.. ESSflI AIIf,PTE ALPH ALPHC M') PT rr RC CD C·L cr'1
.NOH AD 0::8) OK E+06
ADUH t·jfi -l.(H) .-.,. -.1:: ..... 1 · ~:: ';.';- 1. 615 21:"f.~ • 2.4 • ~)272 .2tj9 -. ~~16::!
ADUO tli=t -2 •• )0 -1.78 • I,) 9';' 1.610 2'~~5 • 2.4 • (IU34 .212 -.IZf;"l AD1!1 NR -:3.00 -2. 135 .6',8 1.59:3 2 t3? • 2.4 .01';:3 .071 -. ~J:37 AD1!2 t·ll; 2. (H) 2.62 • ~)t;'" 1.601 "'jt~·'
'- .. I • 2.4 .0482 .4:36 · ~J07 ADU:3 t-ll=t :3. ~ZHZ1 :3.6', • r.)I~I~ 1. 60'~ 2'~G. 2.4 • ~~16:36 .52'3 .024 AD114 HI=t 4. O~) 4.71 • E)',9 1.603 ':!II:' -,
t:.. .' I • 2.4 • (l'~O5 .54? • C12~:1 AD115 NA s. ~:.H) 5.75 • p.;'~E; 1,606 2f~7 • 2.4 · 1~~92 • 56::! • ~~12::: ADlkr5 NI=t 6. 'J0 6.77 • ~;',:3 1.5'34 2';'7. 2.4 .1:317 .5:30 • ~j:;:'3 AD117 t·lI=t ~~. ~~H] 4"' • I · ~;'~'~ 1. 6113 2',7. 2.4 • ~~12:3:3 • :3:::7 -. (144 ADll:3 1-111 o. ~:H) • 6'~ • t::'3'~ 2. 2~)'Z-I 2';iI~5 • ::::. :3 · (1251 . 41 :~: -.04:3
---------_ .. _--------------------------------------------------------------------
1. CS body (pressure)
2. test
3. adaptive/non-adaptive
4. F4 clircraft
-15-
...
? .. ~3 • ESSrlI ADAPTE
NOH AD
AD201 t'lf~ AD2~32 i~ AD203 Hi, AD204 t·lf, AD2(1t:' f~ 1:ID2~)7 Hi, ADzel::: f, AD2"f3 HH AD21~) f~ AD211 t·jf, AD212 H AD213 Hf~ AD214 fl AD215 Nfl AD216 t-lA AD21"? NFl AD218 FI AD219 ~lFl AD220 A AD221 I~ AD222 HI1 AD22~3 t·Hi AD224 11 AD225 HI1 AD226 11 AD22i1 HA AD22::~ A AD22Sl NA AD230 A AD231 HA
.i~ Avion
ALPH ALPHC
;). CH~1 · 12 I). ~H3 · 12 :2. ;:1;) 2.54 4. ~:::1€1 4.9:3 4.0(1 4 '~7
• .'1
6.00 7. 4:~: 6.00 7.41 :3 • ()~:::1 ,-, ...,.~ . ." ... ' .. ' :3 • ~3;J :3.7'6 :3. (10 :3. '?I5 :3. ~)(1 :3. '3(1 :3. ;)~:::1 4.02 3.1i.l1) 4. ~) 1 3.00 4.07 8. ;)0 4.0'3 3. (n) 4. 1'3 3.013 4.2:3 :3 .130 4. 1 ::: :3. (H3 4. 18 :3. ~)i:':1 4.2:3 3. (h) 4. :35 2.013 2.56 :2. ~)I) 2.55 2.80 2.64 2. (H) 2.65 2.1)0 .... -:tIel
~I ,'Ij
2. ;)0 2.77 2.00 2. :32 :2. en) 2.:::4 2.0(1 2. '35
.6% • ~.'O0 • ?132 .6'37 • ?~3(1 .6'j6 • 6'3:3 .69:3 .7'00 · 751 • 755 • 71~~: • 71~7 • 7'3::: • :3109 • !325 .:::44 • :316 • :327 • :342 • :33S~ .7lJfj .701 .751 .755 .:305 • :3P) • :324 • :345 • :352
canard (balance) 116/
PT T'r RC CD CL cr'1 (8) Of( E+(16
1. 611 "'~':I 2. " "" .' '.J .. • (11:31 -. (14'~ · 02'j 1.605 297. 2. I~ • (I H::O -.(14'3 .. (121~ 1.'615 2 t,? . :3 • ~) · e 1:3:3 · 0:3:3 · ~):31 1. 59'~ .,,:,-'
~.' ( .. 2. '3 .0277 .-" .. , C" ~:::1:~::3 • ..::.::..j · 1.607 2'='(:' • 2.9 .0276 .220 • >3:;::::: 1.6136 2970'. 2. " · 13508 • ::::62 037 1.6135 2':'1(' • 2.'~ • ;3474 .. :352 .. ~1:37 1.60a . ., ,:,-, ..... ( . 2. '51 .0217 · 142 032 1.6137 297' • 2. '~ .0216 · 151 • O:~::3 1.695 296. ., .,
.... ' . .:.. • ;:32'28 · 171 · 0::;:3 1.6313 297. :~: .. 1 .0222 · 1 .- .-, O·.j · O:;::~: 1.653 2 1517. .-:. '?
.~ II "- .. tl2:32 · 1:30 • ~3:::::3 1.654 297. 3.3 .'3226 · 174 .13:3:3 1.687 2'~7 • 3.3 1.679 2t~6. :3. :3
" (12:3:3 · 1 :::4 • ;~:3:3 .13:238 · 1 :::6 · 0::::2
1.747 297" 8.5 • (1241 · 195 .. ~~1:32 1.75'~ 2',7. :3.6 .13:230 · 19:3 .0:3:3 1.75'3 2'~7 • :3.5 .0225 · 194 · 13:32 1. 76:3 2';'8. 3.5 .0229 · 1 :37 .. ~3:33 1.763 ;2';t9. :3.5 ,13229 · 1 ,::,.,
" - .13:3:3 1.869 2'~8. :3.8 .0243 · 2;37 · 0:3:2 1.6':):3 21~6. :3 ~ 1 .0176 .0:::0 • ;3::::1 1.695 2'517. .... 1 .:J. .016':J · 07:3 .0:31 1. 716 2'~6. :3.3 .01:35 • ~3:36 .0:31 1. 719 295. :3 11:3 .13 1 :31 .0:34 · ~3:31 1.746 21~6. :3.5 .0185 • ~31~7 .0:34 1.748 295. :3.5 .01 :34 • ~Z1t,:3 · 0:3:3 1.75:3 296. :3.5 · 01:39 · 1 (1~~1 .0:33 1.757 296. :3.6 • \) 1 :36 • ~3t3:3, .~B4 1. :361 295. :3 .. :3 • (11'Z'4 · 1 10 · >3:~:4 -------- .. -.~---------------------~~~---------------------------------------._----
~SS! V!D~T~ ALP:a?ear~T (p:res;;onJ NON AD (8) OK E+06
AD250 NA ~). 00 • 6'~9 1.6';;':3 297. :3. 1 AD251 A (\.00 .. 701 1. 69:3 2',8. :3. 1 AD252 HA 2.01i.l .696 1.698 297. 3.1 AD253 A 2.00 .699 1.694 297. 3.1 AD254 HA 4. i,)(l .697 1.688 2'%. :3. 1, AD255 A 4.00 .700 1.7100 295. 3.1 AD256 t·lA 6. eo .6% 1. 700 2'%. :3. 1 AD257 A 6.0e:l.6"S 1.703 2'~7. :3.1 AD25:3 NA :::.00.6% 1 • 6'?~; 2'?7. :3. 1 AD25'ZI A:::. ~,0 .695 1 • 6,?~3 2'317. :3. 1 AD260 NA 3.010 .698 1.695 297. 3.1 AD261 A 3.00 .700 1.6'?4 295. :3.1 I~D262 NA :3.>;)0 .• 751 1.726 2'%. 3.:3 AD26:~: A ::!, ~j~~1 .757 1.726. 2';"6 •. :3.:3 AD264 t·lA J. 130 .798 1. 745 296. :3.4 AD26:i A :3.0(1.:310 1.7452'%. :3.5 AD267 NA 3.130 .841 1.:::75 295. 3.8 AD268 A 3.00 .876 1.887 296. 3.9 AD269 NA 3.00 .799 1.685 296. 3.3 AD270 A :3. (Hj • :::06 1.689 2'?4. :3.4 AD271 HA 8.121(1.6'37 1.7012'316.3.1 AD272 A:::. el~j .696 1. 68? 2';'5 Q :3. 1
· .. ·----------··-----·--------------------·-::16:..----------.----------------------.. -.. -
.
'~,':"" .
',' ..
TERMS FOR PAGE /16/:
1. canard aircraft (balance)
2. test
3. adaptive/non-adaptive
4. canard aircraft (pressure)
--17-
....
01
• 02t~ i .132·~
· (1:31 • 13:33 • ~1:3:3 • ~137 • 0:37 • 0:32 .03:;; • a:~::~: • ti:3:;:
0·-···· • .:;.:J • 0:33 .033 • 1)32 • 13:32 .0:33 • 0:32 .1333
~ • 138:3 , • 0:34 ) .031 3 • 1~1:31 :- .1331 .j. '~\:31
? J34 :.: .13:3:3 ~1 .03:3 3 • 0:34 ~1 .1334 ---------.
._._-----_ ....
- 17 -LIST OF FIGURE~
F:~ - T2 wind tunnel.
.t~ - CERT T2 test section - C5 body in the test section.
FIG. 3 - C5 model.
FIG. 4 - F4 model • --FJ~ - Mach number distribution along flexible wall with CS.
!J~ - Mach number distribution along flexible walls model •
!}~ - Flexible walls shapes with C5 model •
FIG. 8 - Mach number distribution around C5 model.
FIG. 9 - Mach number distribution around canard model.
!~~ - C5 model Mach number distribution (Me • 0.6, 0.7, 0.84) •
FI~ - C5 model Mach number distribution (Mo > 0.8) •
!l~ - Comparison between TuB and T2, results at Mo ~ 0.7 •
FIQ.:...ll - Comparison between TuB and T2, results at Mo .. 0.84 •.
!!~ - Comparison between. NASA Ames and T2, results at Mo ~ 0.7 •
!~~ - Comparison between NASA Ames and T2, results at Mo C 0.84 •
FI~ - Comparison betwI!en NASA Ames and T2., results at Mo :a 0.95 •
.FI~ - Comparison between NASA Ames and T2,results at Me" 0.97.
FJ:~ - Comparison between calculation and experiment on CS model.
ll_~ - Comparison betwflen calculation and experiment around C5 model
and its sting •
FI!~ - Mach number disl:ribution along flexible walls with F4 model.
~. 21 - F4 model, lift, d)~ag and pitching moment, M '" 0.7.
!f.~ - Canard model IHt coefficient.
!.~~ - Canard model, Hft:, drag and pitching moment. M "'" 0.7.
~;. 24 - Canard model lift coefficinet - MO effect.
~~ - Canard model lift coefficient versus Mo'
m~ - Mach humber distdbution along flexible walls with canard model.
F~;. '27 - Mach number distribution on the canard model (Mo .. 0.7).
!!~~ - Adaptation effect on the canard model Mach number distribution
FIG. 29 ----FIG. 30 ---
(Me a 0.7, a = 0.8°).
Mach number distribution along flexible walls with the canard model.
Flexible wall shapes around tha canard model.
FIC.!..ll - Mach number distribution on the canard model (Me effect).
£!g~ - Nominal and real (balance or forto) angle of attack with the canard
model.
-18-
- 18 -
hester
4~~_: ____ J-------- [~ pressure 80 bar 1 t ~ .~-J
J
FIG 1: T2 WIND TUNNEL - ADAPTIVE WALL TEST SECTION -19-
.. r;:. - 19 -
N2
~==::l:11 !ISlS \ :1":;\1 \ "IS;;;;;
'. ~. ---1430---SCREW EXlERNAL ADJUSTMENT
;
FIG 2: T2 TEST SECTION - C5 MODEL IN THE TEST SECTION' IN
-20-
~ . -...... -,~.-,-.-.. -~-------~ ... _.~-.-~ •• ,. __ ....... _< .. ____ ~_ • ...,._ .. ___ ~_....--.--~~ ____ -~---~- • ,- • I
),,~Wi'>V~?~;~~~'!~~~~~~'W~r;~;mw.m~~·l'r-~
~
i~':;;:
I N I-' I
•
Position of pressure holes
X (mmY 0 6 I 22 I 391 47 I 61 I 77
ellipsoid cylindrical
1/'
-1
-. ~ ';:~~~""·2'<':";~· .::,~",,~ -.-~-.:.. ~ -=-:-:;~~::::~ :~ .. ~~-~ .. :~
~~~~. 1 . ~r:-·"'~··~· .. rk_~""" . \ _. __ ., . ___ ~=~-. .-.;~;: xJ rtF 'Marc .~
lamj~ar turbulent
FTf"' 3' Ph uoncL l.lJ • l>~ '" UL
N o
J
1 1 , 1 , J
.. ......
en o
-E::::.:- ------------.....
t I ----------_. ~. ~ 119.89 J
------------------~ r--.--------1II_-lJIBmm~." .. '
"
.,/-' - .. ...- ,..
FI.G 4: F4 MODEL
-22-
. ,
,J, 'J
.8S
.eo
.75
.. 70
.130
I U « :E
- 22 -
-----------------~
upper wall
. ·lower wall
PLANE WALL
-------,---,----~~------~
A02~ I
AD15
AD31 ~
, X (mm) . ~,!S ·-1 ~~----------~------------~----------~
0 0 0 ~
I
a C)
o ID I
FIG 5: MACH NUMBER DISTRIBUTION ALONG
FLEXIBLE WALLS WITH C5 MODEL
-23-
o o III
•
J2::l ~ , ..
) 15 ~ ..
/
)31
~
)
o ~i~~
./
- 23 -
1.10,. I U « upper wall
1.06 ~ lower wall
1.00
.96
,e
" " " " , : , , , , ,
I , , 1\ I, ;~ ~, 1,1
ll.,. 'II
"
: ~ I I , I
! ~1 I I I I I I I , , I ,
.90+-------------------t------------------~--------------_+~
.96,
•
, I , , , ,
I I ,
\ ' \ , .... J
• 90·, ..... ---------.... ' ------------------+--------------------,4
.95,- A025
.90··~----------------~~----------------·--~-------------------
.9el·
~l""''' _.'~
AD24/'\ ,
.. I....... -.. , . . ',I .. --,'. .. ~' I , .. ,' ..... I ,
.90+------------------+-------
C) C) C) "Ii I
FIG 6:
o o 10 I
o x
MACH NUMBER DISTRIBUTION ALONG FLEXIBLE WALLS WITH C5 MODEL
-24-
....... "" ..... }
(mm) () o LO
' •. &AW# ,a ' .... , ~' %4
- 24 -
r--- -f'. ~. """ (t') tn 9_
CO men 01 01 0.. . . . .
gq.OO 0 X -' 0
U w a R q It 0 0 0 0 0
0 II :::i: ::E: ~ ~ ~
~ I U1 U
I
t , :I:
to I
I-GISt' - I H 3:
l C.I) l.U a. <t :I: tf)
-' -l -' ..J <t <t :3 3: w ..J OJ co Z H X
<1: l.LI -1 ..J 0..
t l.L.
.. l I'-
-~-. C!) 1-1 l l.J...
Ii ... J
-25- ..
I IV 0\ I
•
. SST I U « L
·i ,fj!J ,:;, .. ' ~t;~l.
...... '.
.. '.
i\ <t I
~ fj
FIG 7: FLEXIBLE WALL SHAPES WITH C5 MODEL
"
Mo 0.842 PLANE WALL
•
,'"
" "
"
'. '.
" " , fIJ" •• - •••• '
, ¢ ~.
", ... : .. -.85 '" ..... .....
"
'" .... , .. '
,,#1 .. l." . . -: '"
:'" . / /, ... " : '. . .:'V'.; .' . US' /~'
~ .•.•. '. 0 __ : ___ "'__ _-_Il-<-~~ " ..... :X --- .",,'. --- .' . ... • .......... • . ... ..... j" .+....... •• ~ I •••••
'. • •••••••• • ~ • .,.__ .' • • .. I .' e" ..' .' ......... - . .." . ',' '. • • " ~ •• ' ••••• ;<a- ,..:,.. •• ! .,.. . '. .' . '!iI' .' 'X '. - ..,- • • •••
e
'4" ' •••••• ". -_~: •••• '7"";:'''' ..... '. -:'" ~~ .... "" t .. ··:·.··· . ..... ' ~ .. -~--..., ... .' - .' ~ ./:' .' ... ;..' -,.' .~_. ===g- .... ". .'. -"s' .'
"' .. ;" ,<·f ". ..' .... ..... .... " ". -:*' ... "~ .' .' ~. '. ., .'. -. . . .
" ., ............ - ............... ~ • .' •••• g /' '.' ". .'. .,:' / ...... ". .",- . / ..... ,.:.;.:"" ..
" . . ' ........... .
,,'
" " ,
."
N U\
"/",,,
/. . 831
x (mm) _______
~E---=~=---~-------=-~ ___ ~ o o ru I
d o
FIG 8: MACH NUMBER DISTRIBUTION AROUND C5 MODEL ru
I N r-.
LD m LD 0
d ro
~ II c:s
.-J ~ « 3:
W z « -1 0..
I :
. :'
- 26 -
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f .... :~ .. )~.> ..... / I. • \
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- 27 -
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o 110 ~ -z a H I:::.J CD H ex: IU1 H o c: w co ~ ::::J Z
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-29-
I lit
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t-- <t
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t-
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-30-
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en .::;' 1-..
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- 30 -
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SUo •
~~·--------------+I-------------+I------------~I----------------~mg·a •
pQ
m m m ~ • ...
-31-
c.!l H IJ.
I w tv I
. e
.7( I - I T
Q m Q m ~~ ~~. ~ ~ ~
Q
"" , "1
.....
FIG 13: COMPARlSON BETWEEN TUB A~D T2 RESULTS AT M~O.B4
.ST I I (1 Mo =f17 , ""'_I
v.g. « ~ NASA Ames
.... ____ + h.9.
.s o T2
, .71 I' ,'\.. I "'\ ~ ... ' \ \
\ \ , ,
\ \ I ~
a . .. DB
x/c I . ,
i1 ' • ~5~ Q S N m
a
Q
" FIG 14: COMPARISON BETWEEN NASA AMES AND T2 RESULTS AT M:O.7
s e.g • ...
w --
I w w I
1.1T I i •• u
< 2:
1.f?l+
.9
.s
~ \ \ \ \ \ \ \
(Ii Me =0.84 ., Int v.g.
NASA Ames +- - -th.g.
I .,. 0 T2 MJ::Q.832
m T2 OB43
\ I
2i x/c
• !il
10 N •
.7~~&lf------------------i:------~~--------!!------------------~----~------~~~:J II)
" S L1'l
• ... ~iG 1R: COMPARISON BETWEEN NASA AMES AND. T2 RESULTS AT M:D.B4
W N
1 •
I
. 7(J . f • I, ! . i , ,," . 10 Q III . ~
.1f,J) N 1.1 ~ '9
u •
• S
1.4 T -r
1.2
I ...L. ·u < 4
• •
FIG 15: COMPARISON BETWEEN NASA AMES AND T2 RESULTS AT MfD.84
1--tv1-~- == 0.9 5 -·-i
v.9.\ .... ____ + h.9. NASA Ames
o T2
• ri
w 1. e If::> I
.s
e J I Q .19 ta 10
• ,
... " •
FIG 16: COMPARISON BETWEEN NASA AMES AND T2 RESULTS AT M~O.95
J
• W'4
w w
1 '
. ~
- 34-• 121121 -,;
00 Q)
E <C
« ",.
en « C\J OJ Z l-
i"
e)
C'l
II ci ci 0
>' ,r::: 110 ~
c:)
~: t ' 5,,· t-
I a -< en
I 4-
t-....J ::J en LU a:: C\I t-
o z « en LU ::4
"s; .• <t
« en « z z LU W 3:
....... 0 t-.... W
........ (Il
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en H a: <t
0 sa· 0..
~ 0 c..:l
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~ c.!:l H l.L.
l-I:JYW 1--- I f r
"" N 9 aa II • • • ... ... ...
-35-
I W CI I
~
• e -tl/tl-----.-. " -,;~
In f.il N If)
. III r..
-.
FIG 17: COMPARISON BETWEEN NASA AMES AND T2 RESULTS AT M~O.97 -
0
/'\ 1. aT
I I U
J.
« .9' 2:
.8
I
.7
.6
.5 o T2 exp.
I o4L. calcul -------------------- « ----- ------.-
til lSi
fli
III N
N m T
Ul £'.
0.7
o
x/c -----------f
19 Gil • ....
FIG 18: COMPARISON BETWEEN CALCULATION AND EXPERIMENT ON C5 MODEL
, __ N ___ P/.:_ •.• ..;:; ..... ,:-:'e:t'!".fo'!.t;J "..~.".: ... ~~ .• -..:-*..,,~ .... ~.::.~,.!'!*~~""''"-.:;?:?i>_::~;::_;'0.'.">;'' .. ~f£"__:.~':o<~~r.-,;,~:,;.:_:. :~;::C:i:t~;-:;.;:~,;()-~~~;~~;.:5·~,;,.:-.;;',~~, "-::~~:;:23"~t-'.: .:;s---:;,'>~~:t~.-."' .....
t9 S • r<
w Ul
Q SilCl
:c U « ~
095
...... .........
.. erJ
.. '75
.. 155
- 36 -
---. .,. . --uppal'" wall
lower wall
calculation
--
X (c:m)
.. eCI +-------1--, I o o a
o • o
CD I
o o d "¢ I
o 0 • o d
CD
FIG 19: COMPARISON BETWEEN CALCULATION AND EXPERIMENT
AROUND C5 MODEL AND ITS STING
-37-
-
:m)
:====t o o d
.m
I"MENT'
/
10 I to r-t
II ,..... CD
C::S 3:
I c.. Ol a.
I': Co ::J
0 II
0
::E
- 37 -
r-t r-I CO 3:
C-Q) 3: 0 rl
• I I I I I I \
H8'V1t-J I· I 0 10 ID ,....
to ....-t ....... Cl « I 0 I'
E E
"--'
x
ooe
o
OOGl-
10 OOOi-
10
I w \.0 I
.:>
.6T
-' 0.....0 ."" U ,0
0 • 50§- / /0 .4+ ....
u
/ 03t o.
/ 1M. =0.71 0 .2+ I
• :1 .. o
oj<X (0)
(I) I
o (I) If) m
.6
OSI J .J ...,0.- 0
0 ,.....0 /0
.5 .5 0
I • -4 I a
I I /
.3 0 .3
/ I a .2 a .2 .
• :L t I • :1
0 1'0 Co
o.o-€ • ! 0.0 0 ID 0 III ~
0 0 ~ ~
0
FIG 21: F4 MODEL, LIFT. DRAG AND PITCHING MOMENT. M;O.7
0
o o
~' 0
/0
eM
w ():)
-ri
.,...
" 4
.. ::1
.. 2
.. 1
I j
j .J J :J 1:.: 0. m
. .. J
-. 1
I'
..J tJ
- 39 -
Mo =0.70 ..
€I PLANE WALL
tI ADAPT. WALL
N
FIG 22 CANARD MODEL LIFT COEFFICIENT
-40-
• 40 •
:
°----0--0-0 ____ 0
-0
-41-
, l:
4 tSl M
Q
U
g-
1P -
sa •
r-.. .
f-Z Ud :E. CJ ~
(!) Z r-l :c c...,) r~-I [1-
Cl :2: <C
(.!l .« a: Cl
. llL. H -1
. -1 UJ Cl o :::E
o a: « z <C U
eli C\J
t!J 1-1 L1..
• 41
.4
.J f'. U . 0 110 ~
r: .3 Z
W ::E: 0 :E:
CD Z t-I :r: c.J II l-I-!
./ a. .:2c
a z 4:
/ C..!J ~" <t
/ ex: Cl ADAPT. WALL . .. l-lL. • 1 ED Mo=O.844 I-! .-I
vi! 9 Mo=O.755 .J UJ Mo=O.70 CJ ID 0 ::E:
0 a: <t 0. !3 Z <C c.J
.. (T) C\J
C..!J < 0 ) • Ho- ()( I.J..
• -. 1 -+ + t· --0
51 N "'i' 10 to
FIG 24: CANARD MODEL LIFT COEFFICIENT - M 0 EFFE,CT
-42-
'2L_u~_----o- _o_o_o~~ .. If - -.:::::::----.~ • Q} ----------~~--------~
oc uncorrected
0-0(-3.75°
FIG 25: CANARD MODEL LIFT COEFFICIENT VERSUS Mo
o
') j ''1 C W >
'?'
iJ -f ..J -I J.. .1.. J..l => .. :J (
Ili (\I
c.o H u..
./ I
.715
."10
.se
.715
.'10
.ae
.70
.7!!
• i'O
• f:,a
I U « :t
- 43 -
PLANE WALL
ADAPT. WALL X (mm)
4 +-------------------~--------------.------~--..... ------------..... -~ 0 0 0 ,.. I
0 o () III ,
FIG 26: MACH NUMBER DISTRIBUTION ALONG
FLEXIBLE WALLS WITH CANARD MODEL -44-
o o ID
I .',
• 44 ~
-45-
I ..,. 0'\ I
FIG 27: MACH NUMBER DISTRIBUTluM ON THE CANARD MODEL (M:07)
1.°T I
I~ 2
&--------e PLANE WALL
a m ADAPT 0 WALL
rM-o =0.70 I I ()( = 8° I
.J , , ,X/C, o mom 0 o N m ~ 0 . . . . o ~
FIG 28: ADAPTATION EFFECT ON THE CANARD MODEL MACH
NUMBER DISTRIBUTION (M:07.a =8°)
~ Ul
·510 I U « ~
. /'".\~ '! \. i \ \
::~~"'." , .... \ /.,,1.. /' \ "j \ ,. " .• 4 •• ••• •• " ;;.", .. , _."" .. : ~\ \ ....... , ... \ ........... Y'I!-.- .... ~''t'' 1.'-:: .... ~ \ ,. .. 1.' 1 ! ; " -"fI..' ,,' \., " ... , ••••••••• ,. '" ,"'.I:!% •.•. , I •••• .. \ '"I~ ........, 'o1"'-"j: ......... \~" ..... 'I l 1
'.,. .... :
- 46 -
AD268 '/ J
.Be
AD267
:80
... ______ ) ~pp~r wall
: .. : .. : } lower wa 11
.76
It.. ,~~ ........ ,---, AD263 -,,:. ............. ".... ':. ........ ~ ........ ~.... ....,,'-r-:r."'w::A .. ~~ x. ... ~...... .. ,.'. ;., -.. . "-"" ....... , AD262 ~~~ --' \ "1
AD261 .. 70 .... .: ... "':O~'"':-~.:. " AD260 ~/ .. ~~:;.;.:. .... ~......::=~-~ .... ~ .. ~
o o o ~
I
o o to I
ADAPT. WALL
o
I x (mm)
o o 10
FIG 29: MACH NUMBER DISTRIBUTION ALONG FLEXIBLE WALLS
WITH THE CANARD MODEL -47-
""'~"'g?lfijffC t!I!I!lI9Jl!"~"J. •
- 47 -
~ \\ ,
ADAPt WALL D-I' to X 0 10 '" l' ..J ,.... ,.... CD co UJ Ct . . • . e
Cl1 0 0 0 C;) 0 U II II .II ~ U 0 0 0 0 x :It :Ii: % 0 <:S c:: «
oot>' Z <t U
UJ :J: l-
e z ::J 0
\ c: 0 <:
CJ'J lJ..J 0.. <x: \ " :r: " U1
~
lJJ '\. '\.
Z , -1 < ..J .....I <I: 0.. 3:
oot>'- lJ..J -l CD 1-1 X
'\ LU .....I lJ....
0 ', .. "dZ (11 \\
'\ 009- t!J ~ I--'.\
H .• ~. l1..
0" 0 0 0 oi oi
X (mm) I
0 0 10
. "
-48-
i.6
1.0
I U <! 2.:
- 48 •
IX =3° ~ 01-----6 M 0 =0,7
GJ .
.~ - El
a------e
0.757
OB1.
0.876
iii cockpit
a fuselage
."1
tit \
\ \ .e
• ...G ..',' 'S , . :' '\ \', "s··,···"
. ...." ~:' • Sit' " \:) , '. ' • 'Gl : I "\ \ Q
• I ,,\ .m------a. ... a " 9 -e ~'\ \ ........... ' .... --8 .... __
I r ,.~ --~
// ~ IZl-- -c:J_ Cl ca---- . ~ --c:J -:I ': ~
x/c .6+------4.-'
o o o
Ul (\J
o 10
III
" o o .,..
FIG 31: MACH NUMBER DISTRIBUTIOM ON THE CANARD MODEL (MoEFFEC
FLEXIBLE WALLS WITH CANARD MODEL
-49-
--I o o .
" f"EFFEC!
10
8
..l « w II/
?S
---0-- -
. BAL.ANCE
PRESSURE
6 - - - - - - (X REAL (X NO~1.
(){ REAL
4
z
o ~-----~----~!----~----~'----.~----~I~--~----~ 2 4 68
o
3
--
o (X NOM, = . 0 ----O--G
0--0 ---. G
~_0 __ _
___ - a- ~
__ &V-----<Grr---
----~ -------
2 ----------.---~,-------------------~------0, ~r .0,8 I
FIG 32: NOMINAL AND REAL (BALANCE OR FOTO) ANGLE OF ATTACK CANARD MODEL
-50-
End of Document
