For the case of the F-15 flight tests, boundary layer transition was observed up to Mach numbers of 1.2. For very limited and specific flight conditions, laminar flow existed back to about 20 percent chord on the surface clean up glove. Hot film instrumentation was effective for locating the region of transition. For the F-106 flight tests, transition on the wing or vertical tail generally occurred very near the attachment line. Transition was believed to be caused by either attachment line contamination or strong cross flow development due to the high sweep angles of the test articles. The compressibility analysis showed that cross flow N-factors were in the range of 5 to 12 at transition

Collier, Fayette S., Jr.

Johnson, Joseph B.

Miller, D. S.

Rose, Ollie J.

English

NASA Technical Reports Server

tN90- 12556
SUPERSONIC BOUNDARY-LAYER TRANSITION
ON THE LaRC F-106 AND THE DFRF F-15 AIRCRAFT
Part I: Transition Measurements and Stability Analysis
F. S. Collier, Jr. and J. B. Johrson
Part II: Aerodynamic Predictic.ns
O. J. Rose and D. S. Miller
pRECE.DING PAGE BLANK NOT FiLMeD
997
https://ntrs.nasa.gov/search.jsp?R=19900003242 2020-03-20T00:51:35+00:00Z
PART I: TRANSITION MEASUREMENTS AND STABILITY ANALYSIS
F.S. Collier, Jr.
High Technology Corporation
Hampton, Virginia
J. B. Johnson
NASA Ames-Dryden Flight Research Facility
Moffett Field, California
998
NOTATION
SYMBOLS
F
H,h
M
Cp
Cq
N
x/c
c
U,V
q
P
Re
1
a
A
P
V
m
frequency (Hertz)
altitude (feet)
Mach number
pressure coefficient = P-Poo/qoo
suction coefficient = (PV)w/(PU)oo
N-factor
chord fraction
chord (feet)
velocity (ft/sec)
dynamic pressure (lbf/ft 2)
pressure (Ib/ft 2 )
Reynolds number = UI/v
characteristic length (feet)
angle of attack (degrees)
sweep angle (degrees)
density (slugs/ft 3)
kinematic viscosity (ft2/sec)
SUBSCRIPTS
OD
W
t
cf
free-stream condition
wall condition
transition
crossflow
999
INTRODUCTION
Because of renewed interest in the supersonic flight regime and the recent success
with laminar flow techniques at subsonic speeds, there is a developing interest in
examining laminar flow control for increased fuel efficiency at high speeds.
However, flight data in the area of supersonic boundary-layer transition phenomena
is scarce. In late 1985, a limited window of opportunity of about 6 months existed
where the Langley F-I06 and the Dryden F-15 aircraft were available for flight
tests. Two exploratory supersonic flight tests were conducted to expand upon the
supersonic boundary-layer transition data base. The objectives of these cursory
flight tests were to explore boundary-layer transition measurement techniques and
to identify future experiments. In addition, for future design considerations,
there was a need to obtain transition data to evaluate existing stability theory
criteria at supersonic speeds and to use measured aerodynamic pressure data to
evaluate full potential flow codes in the supersonic regime. This latter point
will be discussed in Part II of this paper.
Primary
• Explore boundary layer transition at supersonic cruise
•Deve]op experience in supersonic transition measurement
techniques
• Identify future experiments
Secondary
• Obtain transition data to evaluate existing staDility theory
criteria at supersonic speeds
• Use measured aerodynamic pressure data to evaluate full potential
flow codes in supersonic regime
lOOO
APPROACH
These exploratory supersonic flight tests were conducted utilizing surface cleanup
gloves on the right wing and vertical tail of the Langley F-106 and the right wing
of the Dryden F-15 test aircraft. Each glove was instrumented with surface
pressure orifices and hot films to obtain measured pressure and transition data
during flight. The measured transition data was correlated with compressible,
linear, boundary-layer stability theory which is considered by many as the
state-of-the-art transition prediction method. The measured pressure data was used
to evaluate a non-linear, inviscid, full potential code which is described in more
detail in Part II of this report.
• Perform explorotory supersonic flight tests using
surfoce cleonup gloves
- F-15
- F-106
• Correlote results with computotions
- COSAL; compressible, lineor boundory loyer stobility theory
- NCOREL; non-lineor, full potential, inviscid code
10Ol
DRYDEN F-15 TEST AIRCRAFT 
The f l i g h t  exper iment  was conducted u s i n g  an F-15 t w i n  engine f i g h t e r  t ype  a i r c r a f t  
w i t h  a wing leading-edge sweep o f  45 degrees. The F-15, n o r m a l l y  used f o r  
p r o p u l s i o n  t e s t s  a t  NASA Ames-Dryden F1 i g h t  Research F a c i l i t y ,  has t h e  a b i l i t y  t o  
reach speeds i n  excess o f  Mach 2. Data f rom p rev ious  f l i g h t  exper iments showed 
t h a t  t h e  F-15 wing produces a p ressure  d i s t r i b u t i o n  which under i d e a l  c o n d i t i o n s  
may y i e l d  smal l  amounts o f  l am ina r  f l ow .  A photograph o f  t h e  F-15 w i t h  t h e  foam 
and f i b e r g l a s s  t e s t  s e c t i o n  on t h e  r i g h t  wing i s  shown i n  t h e  f i g u r e  below. 
ORIGINAL PAGE 
BLACK AND WHITE PHOTOGRAPH 
1002 
INSTRUMENTATION LAYOUT AND TEST CONDITIONS
Two separate instrumentation systems were used for the flight experiment.
Quantities such as Mach number, altitude, and angle of atLack were obtained from
the aircraft's main instrumentation system. The main system employed two absolute
pressure transducers to measure total and static pressure from the flight test
Pitot static probe mounted on a noseboom.
The instrumentation system for the test section consisted ,)f a 32 port electronic
scanivalve, five temperature compensated hot-film anemometers, and an absolute
pressure transducer. The electronic scanivalve was used to measure pressures from
the two rows of 15 flush static pressure orifices and the absolute pressure
transducer measured the pressure on the backside of the scanivalve. The
temperature compensated hot-film anemometers, similar to the system described in
reference 1, were used to measure transition location. Temperature compensated
hot-film anemometers were used because previous experience at supersonic speeds
showed that uncompensated anemometers were sensitive to the local total
temperature, thereby affecting the overall sensitivity of the anemometers between
flight conditions. The temperature compensated anemometers eliminate this problem.
The locations of the hot-film sensors for the first phase cf the flight tests were
5%, 10%, 15%, 20%, and 15% chord. The hot-film sensors were at 1%, 2%, 4%, 10%, and
15% chord for the second phase of flight tests.
Test points were flown at Mach numbers ranging from 0.7 to 1.8 and altitudes
between 20,000 feet and 55,000 feet. Angle of attack ranged from -1 to 10 degrees.
The unit Reynolds number ranged from 1.2 to 4.0 million/f3ot. In order to vary
angle of attack and hold Mach number and altitude constant, constant G-loading
turns were flown.
Flow
C1eonup -_
I. _N.,
oove
°
----J\ 3O7o
Legend
• SUFfOce pressure
orifices
X Hot-film sensors
Test conditions
• Moch no, ronge,,,.,O,70-1,8
• Altitude Fonge,,,,,to 55000 ft
• Re/ft,,,,.,.,,,,,.,1,2-4,0 million
1003
FO/_ AND FIBERGlaSS TEST SECTIOH
The foam and fiberglass test section was placed on the right wing of the F-15 to
eliminate the possible effects of surface imperfections. The test section retained
the existing airfoil shape but added approximately 1/4 inch thickness. The glove
was 4 feet wide and extended past 30% chord. It was constructed using one layer of
unidirectional fiberglass under 1/8 inch thick polyethylene foam covered with four
layers of bidirectional fiberglass. The surface consisted of body filler and
polyester paint. The waviness did not exceed 0,00075 inch/inch.
The test section configuration was changed during the latter part of the flight
experiment. A notch/bump was added to the inboard side of the leading edge of the
test section to eliminate the possible effects of leading-edge contamination.
Four layers of fiberglass
cloth finished with bondo
and polyester paint
Closed ce I 1 po lyuretnone foam 1/4 In,
One layer of glass cloth
Existing wing structure
,2O
z/c 0
-,20 I I I
0 ,20 ,40 ,60
x/c
I
,80
I
i,00
1004
STABILITY ANALYSIS OF THE F-15 SUPERSONIC BOUNDARY-LAYER
TRANSITION FLIGHT EXPERIMENT RESULTS
Shown in the figure below are four test points from the flight test. On the left,
the pressure distributions as a function of chord fraction are presented for Mach
numbers ranging from 0.90 to 1.76, Reynolds numbers ranging from 13 to 18 million,
and angles of attack from about 3 to 9 degrees. The differences in the levels of
the pressure distributions are mainly due to an angle of attack effect. Test
points away from the I-G loading condition were acquired in constant G windup turns
and were considered very nearly steady state. On the right, the corresponding
cross-flow N-factor results are presented as calculated from the compressible linear
stability theory code COSAL. The envelope method was used and the cross-flow
disturbances were assumed to be stationary (zero frequency). As can be seen from
the figure, at M=0.98 and M:l.16, thecross-flow N-factor at transition was 10.5 and
II, respectively. These results are in agreement with previous subsonic transition
correlations that transition should occur where the N-factor is in the range of
9-12. On the other hand, for M:O.90 and M=l.76the cross-flow N-factors were about
6 at transition. There may be several explanations for this result. The above
analysis was conducted for zero frequency disturbances and without the effects of
surface curvature included. It is known that travelling disturbances (non-zero
frequencies)can be more highly amplified, which could result in higher N-factors at
transition. Conversely, convex surface curvature has a stabilizing effect on the
disturbances. It is possible that, with further analysis which takes into account
the aforementioned effects, the N-factors at transition might fall into the
expected range of 9-12.
CD
M = 0,90, Re = 15 x 106 ,
-i,0__,5_° = 3'3° 6 15-(Ncf'__(lO,.:X/Ct)'__= 0'90WM = 0,99, Re = 13 x 10 ,10 ' M = 0,99
0 • O_ Ncf ;,20%)_
Lily M- 1,76, Re : 18 x 106, _-//' M: 1,76
L_ -- Curve f ! t _z .... I Envelope
IiJB(6,2%,' / F = 0 Hz
/ 0 F1]gllt data V .
I I I I I ._-.-.---iJ_.._
0 ,10 ,20 ,30 ,40 0 ,10 ,20 ,30 ,40
x/c x/c
1005
CROSS-FLOW REYNOLDS NUMBER AT TRANSITION
Another method that has been used in the past for correlating boundary-layer
transition data is that of comparing the computed value of the cross-flow Reynolds
number at the point of transition with previous data. Presented in the figure
below is the cross-flow Reynolds number at transition as a function of sweep for a
series of previous incompressible investigations as compiled by Poll (Ref. 2).
The results of the present investigation show good agreement with past results.
400 -
300 -
REcF 200 -
(trons)
100 -
0
10
0 O0
Symbols
c/
13
@
X
[2
I I I I I I I
20 30 40 50 60 70 80
Inci0enc:e
o
7°
8°
3°
0°
0 °
0 °
0 °
0 °
-1 °
-2 °
__o
Source
M = 0,92}
M 0,99 Present
M 1,76
M 0,90
Poll
Gregory
Boltz
AnscomDe
Carlson
Boltz
Boltz
Boltz
Sweep (deg)
1006
F-106 SUPERSONIC BOUNDARY-LAYER TRANSITION
FL I GHT TEST
At Langley Research Center, the F-106 aircraft was utilized to conduct an
exploratory supersonic boundary-layer transition flight experiment. A schematic of
the F-I06 is shown in the figure below. Surface clean-up gloves were mounted on
the right wing which has a sweep of 60 degrees and the vertical tail which has a
sweep of 55 degrees. The gloves were constructed by laminating alternate layers of
fiberglass and epoxy over the existing wing. The final layer was sanded,
primed, and sprayed with lacquer to achieve a very smooth finish. The gloves added
approximately 0.i0 inch thickness to the existing surfaces. During the flight
tests, the Mach number ranged from 0.80 to 1.80, and the altitude varied from 30,000
to 50,000 feet resulting in unit Reynolds numbers in the range of 1.6 to 5.2
million per foot. The angle of attack varied from about 3 degrees up to 14
degrees. Because of fuel and supersonic air space limitations, data had to be
collected in either a slowly varying acceleration or deceleration mode.
F Vertical fin
test Grtic|e
_A_= 55 °
Test conditions
Macn number range-,8 to 1,8
Alt rQnge - up to 50,000 ft
Leadlng edge
test article
A= 60 °
1007
F-106 WING GLOVE 
The surface clean-up glove on the r i g h t  w i n g  of  the F-106 i s  shown i n  the f igure 
below. The glove has a span o f  a b o u t  3 f ee t  a n d  extends beyond 20% chord. Surface 
pressure measurements were made in two rows on the w i n g .  The inboa rd  row consists 
of 18 flush taps cn the glove followed by a Strip-A-Tube pressure bel t  w i t h  12 
t aps .  The o u t b o a r d  row consists of 20 flush o r i f i ce s  on the glove followed by a 10 
t a p  Strip-A-Tube be l t .  The inboard a n d  ou tboa rd  pressure rows were oriented in the 
streamwise direction and were located a t  fractional semispan locations of about 0.5 
and 0 .6 ,  respectively. A s e t  of eight hot-film sensors are mounted on the glove 
for t ransi t ion detection purposes. The h o t - f i l m  sensors were mounted so t h a t  they 
were normal t o  the leading edge. The f i r s t  h o t  film was located approximately 0.5 
inch from the leading edge w i t h  others following a t  in tervals  of  1.5 inches. Since 
attachment l i ne  contamination was anticipated for  t h i s  f l i g h t  t e s t ,  a Gaster bump 
was b u i l t  on to  the leading edge near the inboa rd  edge of  the glove t o  a l l ev ia t e  
t h i s  problem. The best location of  the bump was uncertain because of the u n k n o w n  
migration of  the attachment l ine  t h r o u g h o u t  the f l i g h t  as f l i g h t  conditions change. 
I t s  effectiveness was thus somewhat uncertain. 
ORIGINAL 2?*,;6 
BlAcK AND WHITE PHOTOGRAP4 
1008 
.tlRfe= AGE IS 
F-106 VERTICAL TAIL GLOVE 0g-y 
The s u r f a c e  clean-up g l o v e  f o r  t h e  v e r t i c a l  t a i l  i s  shown i n  t h e  f i g u r e  below. 
There a r e  two s e t s  o f  pressure o r i f i c e s  as on t h e  wing. The i n b o a r d  row c o n s i s t s  
o f  20 f l u s h  s u r f a c e  taps,  and t h e  ou tboard ' row has 16 f l u s h  sur face taps. Each row 
i s  o r i e n t e d  i n  t h e  streamwise d i r e c t i o n .  The inboard  and outboard rows o f  p ressure  
o r i f i c e s  a r e  l o c a t e d  approx imate ly  15 and 3 1  inches f rom t h e  fuselage,  
r e s p e c t i v e l y .  In a d d i t i o n ,  a s e t  o f  e i g h t  h o t - f i l m s  were mounted on t h e  g l o v e  f o r  
t r a n s i t i o n  d e t e c t i o n .  The spacing o f  t h e  h o t  f i l m s  on t h e  v e r t i c a l  t a i l  g l o v e  was 
t h e  same as  t h a t  o f  t h e  h o t  f i l m s  on t h e  wing glove. A Gaster bump was b u i l t  on 
t h e  l e a d i n g  edge near  t h e  inboard  l o c a t i o n  of  t h e  g l o v e  t o  c o n t r o l  at tachment l i n e  
contaminat ion.  On t h e  v e r t i c a l  t a i l ,  t h e  bump c o u l d  be expected t o  be more 
e f f e c t i v e  than on t h e  wing because t h e  at tachment l i n e  l o c a t i o n  (and hence t h e  b e s t  
p o s i t i o n  f o r  t h e  bump) was known. 
1009 
F-106 SUPERSONIC TRANSITION
FLIGHT TEST RESULTS
The table below shows four test conditions where transition was observed downstream
of the attachment line. For all other test points, turbulent flow existed at the
first hot-film gage (0.5% chord). Transition was dominated by attachment line
contamination and rapid cross-flow disturbance growth because of the high sweep
angles. For many of the flight conditions, it seems that the Caster bump, utilized
to control attachment line contamination, may not have worked; but, it has been
effective in other flight tests for controlling attachment line problems and may
have just needed to be "fine tuned". It has been shown in previous studies that
leading-edge suction is very effective in controlling cross-flow disturbance growth
which is present in highly swept applications such as those considered here.
Surface
Wing
A = 60 °
Vertical
A= 55 °
MOo
0,86
1,50
0,91
1,10
Hp (feet)
44000
52000
48500
51000
(_
(degrees)
7,5
5,8
I0,5
7,7
Transition
location
1,5% c
0,5% c
5,0% c
1,5% c
1010
F-106 SUPERSONIC BOUNDARY-LAYER
FLIGHT TEST RESULTS
Presented in the figure below is the effect of suction on the cross-flow N-factor as
predicted by compressible, linear stability theory. The C distribution on the
vertical tail is shown for M=I.72, altitude of 40,000 feet, _nuangle of attack of
about 2 degrees. For the present flights with no suction on the glove, the results
show that the cross-flow N-factor increases very rapidly in the region of the
attachment line. As shown below, Ncf has a value of I0 at 2% chord. With the
suction distribution shown on the right, the cross-flow N-factor is reduced
significantly. The N f has a maximum of about 6 at 5 percent chord. The
disturbances are stabl_ or damped out to about 20 percent chord. This level of
growth is considered to be below the critical value for transition.
Vertical toil
M = i,72 h = 40000 ft, o_ = I,87°
-,5 --
0
Cp
,5
1,0
2O
Ncf 10
0
1,0 --
m
I I
Cq
(xlO00 )
,8
,4
/--Suction off
_Stoble region
  ucT onon" a
,1 ,2
X/C
0
I
.i
X/C
I
,2
1011
F-106 SUPERSONIC LAMINAR-FLOU FLIGHT EXPERIMENT 
The r e s u l t s  o f  these two e x p l o r a t o r y  supersonic  f l i g h t  t e s t s  seem t.o i n d i c a t e  
t h a t  t o  achieve laminar  f l o w  p a s t  t h e  leading-edge wing box c o n s i s t e n t l y ,  s u c t i o n  
must be u t i l i z e d  i n  t h e  lead ing-edge r e g i o n  t o  c o n t r o l  c ross- f low d i s t u r b a n c e  
growth.  Also,  t h e  problem o f  at tachment l i n e  contaminat ion  must be addressed. 
By p r o p e r l y  des ign ing  t h e  l e a d i n g  edge, contaminat ion  can be e l i m i n a t e d .  
Present ly ,  t h e r e  i s  growing i n t e r e s t  i n  conduct ing  a supersonic  f l i g h t  t e s t  w i t h  
s u c t i o n  i n  t h e  leading-edge reg ion .  Shown i n  t h e  f i g u r e  below, i s  an a r t i s t ' s  
r e n d i t i o n  o f  how a t e s t  a r t i c l e  i n c l u d i n g  leading-edge box s u c t i o n  f o l l o w e d  by a 
n a t u r a l  l a m i n a r - f l o w  g l o v e  would appear on t h e  LaRC F-106. The f e a s i b i l i t y  o f  
such a f l i g h t  t e s t  i s  c u r r e n t l y  be ing  s t u d i e d  a t  t h e  LaRC. 
1012 
SI.IN_RY
For the case of the F-15 flight tests, boundary-layer transition was observed up to
Mach numbers of 1.2. For very limited and specific flight conditions, laminar flow
existed back to about 20% chord on the surface clean-up glove. The hot-film
instrumentation was effective for locating the region of transition.
For the F-106 flight tests, transition on the wing or vertical tail generally
occurred very near the attachment line. Transition was believed to be caused by
either attachment line contamination or strong cross-flow development due to the
high sweep angles of the test articles.
The compressible stability analysis showed that cross-flow N-factors were in the
range of 5-12 at transition. Future plans are to analyze the results including the
effects of surface curvature and non-stationary cross-flow disturbances. It was
shown that cross-flow Reynolds numbers at transition correlated well with previous
incompressible results.
F-15 Flight Tests
• Boundary layer transition was observed up to Mocn numbers of 1,2
• Hot film instrumentation worked well for locating the region of
transition
F-106 Flight Tests
• Transition on the wing or vertical toil generally occurred very
near the attachment line
• Transition was believed to be due to either attachment
line contamlnotlon or strong crossflow development due to the
nigh sweep
Stability analysis
• N-factors at transition were in the range of 5-12
• Crossflow Reynolds numbers at transition correlated
well with previous incompressible results
1013
REFERENCES
o
.
Chiles, Harry R.; and Johnson, J. Blair: Development of a
Temperature-Compensated Hot-Film Anemometer System for Boundary-Layer
Transition Detection on High Performance Aircraft. NASA TM 86732, 1985.
Poll, D. I. A.: Some Aspects of the Flow Near a Swept Attachment Line with
Particular Reference to Boundary Layer Transition. Cranfield Institute of
Technology, CoA Report No. 7805, 1978.
1014


Supersonic boundary-layer transition on the LaRC F-106 and the DFRF F-15 aircraft. Part 1: Transition measurements and stability analysis

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