4.

NLR 7301 airfoil

contributed by
National Aerospace Laboratory NLR
Amsterdam, The Netherlands

4.1 Introduction
The NLR 7301 airfoil was selected because it represents the thickest (16.$) of all
supercritical airfoils submitted for inclusion in this data base and appears t o be
close to the limits of useful exploitation of the supercritical shack-free airfoil
concept.
Because of the rather extreme nose radius the airfoil represents probably a hard test
case for Cartesian grid based methods, in particular if also based an the transonic
small perturbation assumption.
The faot that the airfoil represents a rather extreme specimen f thick, aupercritical
8
airfoils is, in the present law Reynolds number teats (= 2 x 10 ), reflected, i.a., in
the typical variations with angle of attack and Maoh number of the aerodynamic coeffi-
oients. It appears that, even at aubcritical conditions, and both with free and fixed
transition, the boundary layer on either the upper surface, or the lower, or both, is
stressed to the limits or beyond. As a result the effects of variations in trsneition
position and transition fixing are rather dramatic. This situation suggests that the
airPoil would be a difficult test case for all methods involving coupled inviscid flow
and boundary layer computations. Note that when transition was free, it occurred generally
thmugh a laminar aeparati~n~hubble which is often reflected in the pressure distribution
(see e.g. fig. 4.10, o = .85 , M = 0.5 - 0.7, lower surface 509b chord). With fixed transit-
ion at 3C$ chord the trip caused generally a local perturbation of the presaure distribut-
ion.
A point of concern (although not a privilege of the present tests) is the amount of wall
interference contained by the test data. The wall interference corrections given in
section 6.2 of the data set have been determined by correlating classical ventilated wall
interferenoe theory with the downwash determined experimentally by means of solid/slotted
wall comparisons and by means of the "brag balance" method (difference between wake drag
and pressure + friction drag). At the time of the preparation of this data set, work was
in progress to determinewall interference from measured static pressure distributions
near the top and bottm wall. The reader is encouraged to watoh the literature for
publication of this work. It is emphasized that the observed difference between potential
flow and experimental design Mach number for shock-free flow for this airfoil of 0.026
(fig. 4.1) is not a measure of the blockage in the testa. The difference can be explained
by the viscous de-cambering near the trailing-edge. The decambering causes a reduction
in circulation and an associated loss in upper surface supervelocity. To restom the
local (shook-free) Mach number distribution the angle of attack and free stream Mach
number have to be increased above the potential flow values.
It is worth mentioning that the tests were done with the specific purpose of verifying
for the first time the aerodpamic characteristics of a aupercritical shock-free airfoil
designed by means of the Boerstcel hodograph method. The airfoil has also been tested
under oscillatory conditions in the same tunnel. At the time of preparation of this data
set a program of high Reynolds number tests in the Lockheed Georgia Compressible Flow
Facility was partially co leted andtestsin the NASA Ames 11 ft x 11 Ft tunnel (steady
and unsteady, Re ~ 1 x5loz? were about to be started.
4.2 DATA SET.
1. Airfoil
1.1. A i r f o i l designation NLR 7301 ( a l s o NLR RT 7310810)
1.2. Type of a j r f o i l t h i c k , aft-loaded, shook-free s u p e r c r i t i c a l 5
designed by means of Boerstoel
hodograph method

1.2.1. a i r f o i l geometry see f i g . 4.1 and t a b l e 4.1

nose r a d i u s Roc"%
maximum t h i c h e s s t/0 = 16.3%
base t h i c h e s s eero

1.2.2. design condition p o t e n t i a l flow (hadograph theory):

M = 0.721
c1 = 0.60
experiment ( f r e e t r a n s i t i o n , NLR P i l o t ~ u n n e l ) :
Mt = 0.747, clLs 0.45
design pressure d i s t r i b u t i o n see f i g . 4.1 :t a b l e 4.1
1.3. Additional remarks design method described i n r e f . 1
1.4. Reference8 on a i r f o i l none

2. Model ~ e o m e t q

2.1. Chord length

2.2. span
2.3. Actual model 00-ordinates and
accuracy see t a b l e 4.2 and f i g . 4.2
2.4. Madmum t h i c h e s s t/c = 1 6 . 9
2.5. Base t h i c h e s s 0.1% chord
2.6. Additional remarks f i n i t e trailing-edge (base) t h i c h e s s was
obtained by cutting-off t h e o r e t i c a l a i r f o i l
a t 9 8 . 9 chord
2.7. References on model none

3. Wind t m e l

3.1. Designation NLR P i l o t t m e l

3.2. Type of t-el continuous, closed c i r c u i t
3.2.1. s t a g n a t i o n pressure atmospheric
3.2.2. s t a g n a t i o n temperature +
313 - 1 K
3.2.3. humidity/dew point v a r i e s with atmospheric condition
( s t a g n a t i o n temperature chosen such t h a t
condensation i s avoided)
3.3. Test s e c t i o n see f i g . 4.3
3.3.1. dimensions rectangular
height0.55 m, width 0.42 m
3.3.2. type a f w a l l s 1046 s l o t t e d top and bottom w a l l s ,
solid side walls
separate t o p and bottom plenums
3.4. Flow f i e l d (empty t e s t s e c t i o n )
3.4.1. reference s t a t i c pressure taken a t aide wall 3.6 chords upstream of
model
3.4.2. flow angularity upwash Ao = 0.12' (+ 0.03')
( w i t h r e s ~ e c tt o tunnel reference plane)
3.4.3. Mach number d i s t r i b u t i o n see f i g . 4.4a
3.4.4. pressure gradient see f i g . 4.4b
3.4.5. turbulenoe/noise l e v e l see f i g . 4.5 and r e f . 4
3.4.6. s i d e wall boundary l a y e r t h i c h e s s 1% of t e s t s e c t i o n semi-width,
no s p e c i a l treatment
3.5. Additional remarks f o r two-dimensionality of t h e flow see r e f . 3
3.6. References on wind tunnel ref. 2

4. Tests
4.1. Type of measurements surfaoe pressures ( l i f t , p i t c h i n g moment)
wake p i t o t pressures ( d r a g )
surface flow v i s u a l i z a t i o n
flow f i e l d v i s u a l i z a t i o n
4.2. ~unnel/model dimensions
4.2.1.
4.2.2. /
height chord r a t i o
width chord r a t i o
4.3. Flaw oanditians included i n present
d a t a base
 4.3.1. angle of a t t a c k -4' t o + 4' f o r Mt = 0.747
4.3.2. Mach number 0.30 t o 0.85 f o r at = 0.85'
6
4.3.3. Reynolds number about 2 x 10 ( s e e f i g . 4.6)
4.3.4. transition f r e e and f i x e d
- p o s i t i o n of f r e e t r a n s i t i o n see f i g . 4.7
- transition fixing s i z e 130 (90-106p) b a l l o t i n i ( g l a s s beads)
bands of &m width a t 3046 chord
on upper and lower surface
4.3.5. temperature equilibrium Yes
4.4. Additional remarks 1 ) without boundary l a y e r t r i p t r a n s i t i o n
oocured generally through a laminar
separation bubble.
2) t h e r e a n i n d i c a t i o n s f o r i n c i p i e n t r e a r
separation a t a l l flaw conditions eltoept
a t low c and arolmd t h e design condition
1
with free transition
4.5. References an t e a t s ref. 5

5. Instrumentation
5.1. Surface pressure measurements

5.1.1. pressure holes

-size diameter 0.25 mm; depth lmm
--
spanwise s t a t i o n ( s )
chordwise p o s i t i o n s
staggered (+_20 mm) around c e n t r e l i n e
Bee 4.3 and fig. 4.2a
5.1.2. type of transduoere and one 2 7.5 p s i and two + 5pei Statham
scanning devices d i f f e r e n t i a l Dressure transducers + 48 s t e m
Scaniwrlvea;
reference pressure p-measured with C.E.C.
1 5 p s i absolute prsasure transducer
(aocuraoy 2 0.05 $)
5.1.3. other no
5.2. Wake measurements
5.2.1. type/siee of instrument(6) wake rake ( f i g . 4.3);
69 tubes, spacing according t o t a b l e 4.4
tube diameter o u t e r / i m e r : 1.0/0.7 ma
5.2.2. atreamwise position(^) 0.8 chords downstream of t r a i l i n g edge
5.2.3. type of transducers and two 2 2.5 p a i Statham d i f f e r e n t i a l pressure
scanning devices transducers + 48 s t e p s Scanivelves;
reference pressure p ,
meaeured with C.E.C.
1 5 p s i absolute presaure transducer (accu-
racy 2 0.05 %)
5.3. Boundary l a y e r measurements no
5.3.1. type/size of instruments
5.3.2. l o c a t i o n s
5.3.3. type of transducers and
scanning devices
5.4. Skin f r i c t i o n measurements
5.4.1. type/aize of instruments
5.4.2. l o c a t i o n s
5.4.3. type of transducer
5.5. Flow v i s u a l i s a t i o n
5.5.1. flow f i e l d shadargmph picture8
5.5.2. surface flow d e t e c t i o n of t r a n s i r i o n position b
eublimation Technique (acenaphteney
5.6. o t h e r no
5.7. Additional remarks length of pressure tubee 4m;
scanning r a t e 2 preesures/sec.
5.8. References on instrumentation none
6.1. Accuracy (wall i n t e r f e r e n c e excluded)
6.1.1. angle of a t t a c k s e t t i n g
6.1.2. f r e e stream Mach number:
- setting
- v a r i a t i o n during one
2 0.001
pressure soan
6.1.3. pressure c o e f f i c i e n t s ACD =*0.002 t o 0.02 depending an looal
pressure l e v e l and d p a m i c pressure
6.1.4. aerodynamic c o e f f i c i e n t s "nk"0wn
6.1.5. boundary l a y e r q u a n t i t i e s n.a.
6.1.6. repeatability Aclwt 0.004; Acd%+_0.0005; Ac
m
*+_
0.001
6.1.7. remarks none
6.2. Wall i n t e r f e r e n c e corrections
( i n d i c a t e estimated accuracy)
6.2.1. angle of a t t a c k Aa = -1.4 x c 1
+ 0.56 ( c +0.25cl)
~ / m(degrees)
6.2.3. streamline curvature

6.2.4. other
6.2.5. remarks wall interference i s presently being
reassessed
6.2.6. refemnces on w a l l i n t e r f e r e n c e ref. 6
corrections
6.3. Presentation of data
6.3.1. aercdpamic c o e f f i c i e n t s fige. 4.8, 4.9 ,
t a b l e 4.5
6.3.2. surface pressures t a b l e 4.5 ; fig. 4.10
(table4.5 includes w a k e rake gressures)
6.3.3. boundary l a y e r q u a n t i t i e s -
6.3.4. wall i n t e r f e r e n c e corrections tabulated d a t a f o r a , cl and c a r e presented
m
included?
with and without c c r r e c t i m s f o r downwash and
streamline c-ture. No blockage corrections
No corrections on c and C Figures present
d P'
only uncorrected values
6.3.5. corrections f o r model deflection no
6.3.6. Empty t e s t section c a l i b r a t i o n
taken i n t o account? no
6.3.7. other corrections included? no
6.3.8. a d d i t i o n a l remarks f i x e d t r a n s i t i o n surface pressure data a r e
a f f e c t e d by l o c a l disturbances due t o
t r a n s i t i o n band ( i n p a r t i c u l a r hales number
15, 47 and 48)
6
6.4. Were t e n t c a r r i e d out i n d i f f e r e n t 1) Tests a t various Reynolds numbers (3-30110 )
f a c i l i t i e s an t h e current a e r o f c i l ? conducted i n Lookheed Ga. Compressible
If 8 0 , what f a c i l i t i e s . Are data Flow F a c i l i t y . Not included i n present
included i n t h e present data base? d a t a base
2) Unsteaqy ( o s c i l l a t i n g a i r f o i l ) t e s t s on
other model i n same tunnel. Not included
i n data base.
3) U n s t e a q ( o s c i l l a t i n g a i r f o i l ) t e s t 6 a t
15 x 10 Re number planned i n 2-a t e s t
set-up of NASA Ames 11 x 11 f o o t Tunnel.
Not included i n data base
6.5. To be contacted f o r f u r t h e r J. Zwaaneveld
i n f a m a t i o n on t e s t s National Aerospace Laboratory NLR
Anthony Fokkemeg 2
Amsterdam 1017
7. References

1. J.W. Boerstael "Transonic shock-free a e r o f o i l design by

G.H. Huizing a n a l y t i c hodograph methods"
NLR MP 73023 U
Also AIAA Paper 74-539
2. J. Zwaaneveld P r i n c i p a l Data of t h e NLL P i l o t Tunnel
Report MP. 185
3. H.A. Dambrink I n v e s t i g a t i o n of t h e 2-dimensionality of
t h e flow around a p r o f i l e i n t h e NLR
0.55 x 0.42 m2 t r a n s o n i c wind tunnel
NLR Memorandum AC-72-018
4. R. R O E S N o i ~ eenvironment i n t h e NLR t r a n s o n i c wind
P. Rohne tunnel HST
NLR TR 74128 U
5. J. Zwaaneveld Aerodpamic c h a r a c t e r i s t i c s of t h e s u p e r
c r i t i c a l shook-free a i r f o i l s e c t i o n NLR 7301
6 . J. Smith Valuee of w a l l i n t e r f e r e n c e c o r r e c t i o n s
f o r t h e NLR P i l o t Tunnel with I@ open t e s t
section
NLR Memorandum AC-74-01

8. L i s t of Symbols
8.1. used i n t e x t and f i g u r e s
C pressure o o e f f i c i e n t
!c c r i t i c a l pressure c o e f f i c i e n t
Y
c a i r f o i l chord length
0.4
A
.
&reg c o e f f i c i e n t
lift c o e f f i c i e n t
c p i t c h i n g moment c o e f f i c i e n t ( w i t h respect t o .25c)
m
M f r e e stream Mach number
p_ f r e e stream s t a t i c preseure
f r e e stream stagnation pressure
3 m a m i c pressure
Reynolds number
Reo
R*
t
- leading edge r a d i u s
a i r f o i l maximum t h i c h e s s
X,Z a i r f o i l coordinate system
x+, z+ windtunnel coordinate system
" "
01 angle of a t t a c k
subscript
t r e f e r s t o uncorrected values

8.2. used i n d a t a t a b l e s

ALWA
ALPHAT t* ( w i t h reepect t o tunnel reference plane)
CDPB pressure drag c o e f f i c i e n t , uncorrected
CDP pressure drag c o e f f i c i e n t , corrected f o r w a l l i n t e r f e r e n c e
CLB
clt
o1 ,
corrected f o r w a l l i n t e r f e r e n c e
C

CM c , corrected f o r wall i n t e r f e r e n c e
CPIB C; (uncorrected)
t o t a l head d e f i c i t pressure c o e f f i c i e n t i n wake
MAB f r e e stream Mach n m b e r , unc r r e c t e d
PI l o c a l s t a t i c pressure ( ~ / m) 3
m stagnation pressure (kgf/m2)
f r e e stream dynamic pressure (kgf/m2)
Rec
curvature
surface slope