A transonic wing, the DFVLR-F4 was designed and tested as a model in European transonic wind tunnels and was found to give performance improvements over conventional wings. One reason for the improvement was the reduction of compression shocks in the transonic region as the result of improved wing design

Redeker, G.

Schmidt, N.

English

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NASA TECHNICAL MEMORANDUM NASA TM-75752
NASA-TM-75752 19800013773
TRANSSONIC WING DFVLR-F4 AS EUROPEAN TEST MODEL I
G. Redeker and N. Schmidt -i
- !
Translation of "Transsonischer Tragfluegel DFVLR-F4
als europaeisches Testmodell," DPVLR-Nachrichten, June i
1979, pp. 36-39 i
FEB! 11990
LANGLEY RESEARCH CENTER
LIBRARY,NASA
"_MPTON, VIRg'*"
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
WASHINGTON, D.C. 20546 JANUARY 1980
https://ntrs.nasa.gov/search.jsp?R=19800013773 2020-03-21T18:07:24+00:00Z
STANDARD TITLE PAGE I
,._o_ NTM_7 57 52 J 2. Go .... ment Acces,Jon No. J 3. Recipient's Catalog No. _i
)•
4. Title and Subtitle 5. oort D.._
• !_.f
.rRANSSONIC WING DFVLR-F4 AS EUROPEAN January 1980
TESTMODEL 6.PerformingOrganizationCode
7. Author(s) 8. Performing Organization Report No.
Dr.-Ing. Guenter Redeker,
X 10. Work Unit No.
Dr.-Ing. Norbert Schmidt
-- 11. Contract orGrantNo.
9. PerformingOrganizationName andA4dress NASW-3199
Leo KannerAssociates |13. TypeofR,DortariaPer+odCo,area
Redwood City, California94063 I Translation
|2. Sponsoring Agency Name and Add,ess
_ational Aeronautics and Space Adminis-
tration, Washington, D.C. 20546 14.SponsoringAgencyCode
15. Supplementary Notes
Translation of "Transsonischer Tragfluegel DFVLR-F4
als europaeisches Testmodell", DFVLR-Nachrichten_ June .....
1979, PP. 36-39
16. Abstract
Commercial aircraft of the next generations flying
at high subsonic speeds in the neighborhood of 0.8 Mach
will have transsonic wings. Such a wing, the DFVLR-F4
wasdesigned and tested as a model in European tr_nssonic
wind tunnels and was found to give performance improvements i !
over conventional wings. One of the reasons for the ii
improvement was the reduction or eliminat%on of compression i
shocks in the transsonic region as the result of improved j
wing design.
17. Key Words (Selected by Author(s)) 18. Distribution Statement
Unclassified-Unlimited i
19. SecuritYunclassifiedClassif.(of this report) I _"UnclassifiedSeCurityCla s f. (of this page) 21. No. of Pages 22. Price
!
+
i i
NASA-HQ i
+
TRANSSONIC WING DFVLR-F4 AS EUROPEAN TEST MODEL
Dr.-Ing. Guenter Redeker, Dr.-Ing. Norbert Schmidt
Institut fuer Entwurfs-Aerodynamik der DFVLR, Braunschweig
GARTEur (Group for Aeronautical Research and Technology in Europe)
is an institute with several expert groups in the area of aeronautics
in Europe. In group 6 "Supercritical Wings" the Federal Republic of
Germany, Great Britain, France, and the Netherlands, are working on
the technology of the transsonic wing whose function is to make sub-
sonic travel more economical by savings of fuel, by increases in range,
and by increases in pay load. At the last meeting of the GARTEur
group 6 in November 1978 the decision was made to test the wing
DFVLR-F4 as a wind tunnel model in several European wind tunnels.
At the same time the GARTEur 6 partners are carrying out aerodynamic
calculations on this wing shape_ valuable results in the field of com-
puters as well as in the area of wind tunnel technology are expected
herefrom. The following report will briefly explain the technology
of the transsonic wing and will report the efforts of the Institute
for Aerodynamic Design which, within the framework of the civilian
component program (ZKP) and in the program "airplanes" of the Research
and Development Program of the DFVLR (German Research and Test Institute
for Air and Space Travel) led to the design of the transsonic wing
DFVLR-F4.
Technology of the Transsonic Wing
Commercial aircraft of the next generation flying at high sub-
sonic Mach numbers of Ma=0.8 (e.g., the Airbus 310, figure 1:)will
have transsonic wings. Here a considerable fraction of the lift at
the wing will be achieved by reduced pressures on the upper surface
of the wing which are brought about by local supersonic fields. The
_otential_or_nc_eas_ng_lift with supersonic fields is generally known_
however, a compression shock develops in conventional wings on the upper
1
surface during transition from supersonic to subsonic conditions
(figure 2, left). This often leads to flow separation and increased
drag at the wing so that an economical and safe flight is not possible
with conventional wings in the transsonic region.
Figuge 1_ Typical commercial airplane for high subsonic speeds
However, with the new wing shapes based on transsonic wing tech-
nology it has become possible to reduce extensively these unfavorable
effects. Here pressure distributions at the wing are generated which
have large supersonic contribut£D_s and which generally experience no
or only very weak compression shocks (figure 2, right) and which thus
produce a sensibly economic flight. These improvements in the aero-
dynamic characteristics can be utilized in various ways to increase
the economics of commercial aircraft as can be seen schematically in
figures 3 and 4.
Figure 3 shows that, by keeping the basic wing cross section and
- _the flight_Machn,_mber _onstant__epback € _nd_flight_Mach
number Ma constant), the wing thickness can be increased, for instance,
from a current _ =10.5% to _ =13% which results in a decrease in wing
weight, resp. an increase in pay load; or if the wing thickness 6 is
kept constant, the flight _4ach number Ma can be increased from, e.g.,
a current Ma=0.78 to Ma=0.83. This produces an increase in transport
performance.
s
1 2 1 - conventional profile
., @ konventionellesProfil @ frQnssonischesProf,l 2 - transsonicprofile
3 _I 3 - compression shockVerdichtu6ngsstol_ -4....supersonicz oneh 5 5 weak shock
scwocher 6 flow separationIx, _<'_. \\.1 Ablssung /x\zone'_\_ gtoA -
_ _ 7 - supersonic portion
__ 8 - local sonic velocity
_ 9 - local wing depth
•_ _,, C_ [ortl,ch SChallge. _:
_ '__,_,_ _ I0- local pressure value
io_ io
/ortliche Profilhefe
Figure 2: Flow field and pressure distribution on conventional and on
-- "i_t _ach numbe_-]the transsonic wing. sweepback
m 074 constant _ constant I
Fldgelpfeilung _ _0 lZ " I
cD konstant : (D Flugmachzahl _ I
0.13 i _ i0, 3 _. Aonstont
"_ BQS,S Airbus 11 .! •O.75 .: _ O.80 0.85 I ' 30 ° 20 ° 10°
_. Flugmachzahl MQ I _ Flugelpfeilung p" (D
' flight.[.4achi'numbe_ _ wing sweepback
Figures 3 and 4: Evaluation of the technological gains resulting from -
the use of the transsonic wing (left, for unchanged wing cross section;
right, for constant flight Mach number).
An additional possibility is shown in figure 4. By keeping the
wing thickness 6 constant the additional aerodynamic potential can be
utilized to reduce the wing sweepback from, e.g., a current _ =30° to
=lO °. This again produces constructional advantages for the wing which,
because of the shortened wing strut, lead to a weight saving. The
question as to which of the possibilities or combinations thereof dis-
cussed here should be used in an individual case depends on the demands
made of the aircraft.
3
Civilian Component-Program and Wing Design of the DFVLR
In 1975 the ZKP "Wing section" panel of the
"20
,cp I BMFT was initiated for the purpose of safety and1.g
1.2 _ _ _ further development of the transsonic wing and of
0_8 decreasing the development risks in new airplane
"__--- projects; this panel, for one, has the goal of con-
0% _i_ _ structing an experimental carrier with a transsonic
Y "\_._ wing for the large transsonic wind tunnel $1 of the
m _
__[]Mo=073 _ 0NERA in Modane, South France, and for another,
-0.8 j[Rej=107jj 1 attempts to improve the aerodynamic development
-1"200.2o% 0_60.8x_0 processes. In this program considerable contributions
T were made not only by the German aircraft industry
which, in the course of their efforts, leaned heavily
Q Profil DFVLR - 48080
on the Airbus variation 310, but also by the DFVLR
-_-Co=0.5__ within the framework of its FuE program "airplanes".
For this purpose the Institute for Aerodynamic
(_.Profil DFWLR- R3 Design, in addition to a continued development of
aerodynamic design methods in the years 1975 to
1978 also produced the design of a transsonic wing
QProfil DFVLR - R4 for a demonstration airplane similar to the Airbus
and tested it with good results in the transsonic
wind tunnel of the DFVLR in Goettingen.
Figure 5: Development
of the transsonic wing
DFVLR-F4.
Transsonic Wing DFVLR-F4
For the development of a transsonic wing one must proceed in steps
so that-ini%-ially the wang cross section for a new alrplane does not
differ basically from that of today!s planes (compare figure 1). It will
be a swept-back wing of high aspect ratio. However, the wing shape will
change decisively since "supercritical" or transsonic wing shapes with
4
a relatively blunt nose and a flattened upper wing surface and increased
curvature in the rear wing section will be employed here. Designs of
commercial aircraft are mostly directed toward the cruise condition,
i.e., the requirement exists to transport a pay load over a given dis-
tance at a given flight altitude with a given flight speed. If one
starts from the premise that the size and the cross section of the wing
were optimized by preliminary tests of wing loading, sweepback angle,
aspect ratio, and taper with respect to wing weight and fuel costs, then
a design criterion for the lift coefficient and the flight Mach number
of the wing can be derived.
For airplanes with largeaspect ratios A =i0 under consideration
here, the pressure distribution at the wing is determined, size wise,
by the cross-sectional shape of the wing in the direction of the flight.
Thus first a wing shape, always depending heavily on the demands made
of the wing, is developed so that first the complex structure of a three-
dimensional wing can be expressed as a two-dimensional problem.
Figure 5 shows the development of a transsonic wing shape. Starting
with a profile design DFVLR 48080 (case A) with an average lift coeffi-
cient of ca=0.50, a profile (cases B and C) is developed, by appropriate
modifications, which satisfies the design requirements for the wing.
Here the lift coefficient of Ca=0.50 is increased stepwise to Ca=0.65
for the case of profile C without noticeable deterioration of the
profile characteristics. This is achieved, as shown by the pressure
distribution c along the profile depth x/l, by an expansion of the shock-P
free supersonic field and by a change of the pressure distribution on
the lower wing surface near the trailing edge. By means of special shapes
in this region, pressure differences between the upper and lower surfaces
are achieved near the trailing section which contribute in great measure
to the lift and which are known as "rear loading". The decision as to L
whether a profile is suited for the given job is made after extensive
_sign-calcula_ions-with-a-high-speed_eomputer usin_various incident-
flow conditions and by experimental check tests in the wind tunnel.
5
Figure 6 shows the results of such a pressure-distribution test made
on the profile DFVLR-R4 in the transsonic wind tunnel of the DFVLR in
Braunschweig. For the profile incident-flow Mach number of Ma- 0.73,
reduced because of the sweepback effect, the pressure distribution
along the profile depth x/l is shown for two lift coefficients c La
and is compared with the design calculations. Here it is shown clearly
that, on the one hand, the measurements confirm the desired pressure
distributions, but that, on the other hand, the design calculations
also give very reliable results. The profile DFVLR-R4, shown here,
was used as the base profile for the wing design.
_.e-- c_--o.eo -J_o--o.77[-- m_orilDFVL_-R_
-o_ calculation !
Rechnung Re--107I
I I "° MessungTW8Re=6.106
-12o 0.2o%o.6o.8x3oo.2o%o.eo.axl.Omeasurement
1 I
.... . ._ -_
Figure 6. Experimentally determined pressure distributions at the wing
DFVLR-R4.
The design of the wing was carried out under the following
points of view:
- minimal induced drag through elliptical distribution of lift along
the span width_
- utilization of the good properties of the profile DFVLR-R4through
realization of the profile pressure distribution, also on the wing
- inclusion of airplane fuselage in the wing design, i
With-the-aid-of--a-turbulence-ladder process-in-whi_h-thewing is
composed, for theoretical potential calculations, of many horseshoe-
shaped vorices arranged in ladder form, a preliminary design was first
7
I
6 f
made which produced information with regard to the distribution of
distortions along the span width, i.e., the mutual twisting of the
individual wing cross sections. After the extent of this distortion
distribution had been determined, the wing contour was generated by
building up the afore-described profile DFVLR-R4 into four definition
sections of the basic wing cross section.
-" layout ....
Auslegung
[ d/I:0122 EF=.=-u,J j ._::_-05°1 ;Maoo-0 785
cA=0.5 d/I--0122, E=09 °
..%:0 o I " , -F "_
'
1.o x/l
Figure 7i Construction of the wing O T,R-F4
Figure 7 shows several profile sections in a perspective drawing
of the wing where the four definition sections are specially marked by
the local thickness 6 =d/l and by the distortion angle aF. Between
these sections the wing contour was established by linear fairing. In
the area of the transition from fuselage to wing the profile-shaping
of the wing was improved by iterative after-calculations of the wing
using a three-dimensional process for calculating the wing-fuselage
layout in the transsonic range. The wing DFVLR-F4 thus developed was
also tested by extensive calculations in the off-design region, i.e.,
for flight numbers and angles of attack deviating from those specified.
An example of a pressure distribution, obtained by calculations, for
an initial cruise condition at a flight Mach number of Ma=0.785 and
a lift coefficient of CA=0.50 is shm_n in figure 8. This shows clearly
that a local supersonic field with shock-free recompression exists
along the entire span width.
The wing designed here was fabricated as a half-model of a wing-
fuselage layout and was tested in the transsonic wind tunnel in
Goettinge_. The model (figure 9) has a half-span width of s= 587.77 mm
7
and a fuselage length of L= 1192 mm. Pressure tap holes for the measure-
ment of the static pressure along the contour are drilled in 36 places
each at each of the five span width sections 1 to 5. An example of such
a pressure distribution measurement is shown in figure lO. _he pressure
distribution along the local wing depth is plotted, for each of the five
span widths, in the perspective wing cross section drawings. It can be
seen that for the incident-flow conditions Ma= 0.785 and CA= 0.5 there
appears, even experimentally, an extended local supersonic field which
exhibits only weak compression shocks.
layout
Ausle_gu___ng_ I
I II %
>
Figure 8: Computationally determined pressure distribution at the
DFVLR-F4 wing at the initial conditions.
princ, data of model
wzng area A summarizing calculation of the wingspan--width- .....
aspect ratio DFVLR-F4 can be made with the aid of figure II ;
ref.--wing-r-:b._ _-
_ _.,._ In a lift coefficient-Mach number diagram lines
•. _' deptht A •95
: _ taper ,,_:.'_ of constant aerodynamic efficiency are plotted
._-o._
; ' sweepback _..2_,o with the parameter (CA/CW) Ma= constant which
I :i"_osor_o,, ....._ reSa bes lift, drag, and flight Mach number to one
? i--__ another. The higher the plotted value, the better
..... _ _ i _! is_he-aerodynar_ciency_f__the wang ....In the__
-----58z7o_...._ region around r_a= 0.78 and CA= 0.5 to 0.6 one can
,!_- , detect a maximum which includes the design value.
" Thus the experimental testing of this wing shows
press. distrib, that the highest aerodynamic efficiency is_ mea . cro s
sections
Figure 9" View of the wind tunnel model8
attained in the range of the design requirements. Also plotted in
figure 11 is the curve showing the increase in drag which limits, in
an economic sense, flight at still higher Mach numbers
,°11DF VLR - F 4 co cp
Cp . I 0 p._ It--- p'-_.'_ /
-1.o _o,_,,_ Ik, / "
-1,o _ -) _-o.,,2
. _ _ =0.4o_neasuremez[t
,,_ k,_ _/ Messung
, _ r_= O,238 G5 ttingen
'_ Re = 2.106
c,4 = 0.513
"-- ® x/It" -0. 131l CMN2S=
. . . . . .. ,_
Figure 10: Experimentally determined pressure distribution at the
DFVLR-F4 wing at the initial conditions.
These results clearly point out that the
I...,,"o' wing DFVLR-F4 satisfies th sign requirements
io8
........ and that, on hand of the existing knowledge of the
--
0, _ transsonic wing, improvements are possible in the
°' _z_ aerodynamics of today's commercial aircraft. For
°" the wing DFVLR-F4 the potential of the transsonic
_' was utilized in a direction whereby, compared
_, _ ..... I_ rise to the Airbus standard, the average wing thickness
• _gn,, r_qu: was increased from 6 = 10.5% to 12.5% and the wing
.... o, o. _8 ,°_ _, sweepback was decreased from _ = 300 to _ = 27° •
Figure 11_ Measured performance
values of the aerodynamic quality
of the DFVLR-F4 wing.
_ 9
Outlook
Based on the good results obtained and on the fact that a wing
of the latest technology was developed here, the GARTEur group
°'supercritical wings" decided to use the DFVLR-F4 wing as a test model.
This wing will be tested during 1979/1980 in the transsonic wind
tunnels of the ONERA (France), the NLR (Netherlands), and the RAE
(England) as a full model. Valuable information with regard to wind
tunnel corrections are expected from this. Theoretical calculations
for this wing, made concurrently, are expected to give an overview of the
state of the aerodynamic, transsonic computer methods in Europe.
1D


Transonic wing DFVLR-F4 as European test model

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