The National Transonic Facility (NTF) was operated to design condition of 120 million Reynolds number at a Mach number of 1.0. All systems were checked out except plenum isolation valves; modifications are being made to heaters on the actuators. Initial steady-state calibration indicates excellent steady flow characteristics. The first test of the Pathfinder 1 model indicated significant Reynolds number effects. Some effect of temperature on instrumentation were obtained. The cause of these effects is being evaluated

Fuller, Dennis E.

Mckinney, Linwood W.

English

NASA Technical Reports Server

?_88-14941
PRELIMINARY CALIBRATION AND TEST RESULTS
FROM THE NATIONAL TRANSONIC FACILITY
_/_--- _
Linwood W. McKinney and Dennis E. Fuller
NASA Langley Research Center
Hampton, Virginia
311
https://ntrs.nasa.gov/search.jsp?R=19880005559 2020-03-20T07:44:36+00:00Z
IGHTSOF NTFOPERATINGEXPERIENCE
The highlights of the National Transonic Facility (NTF) operating
experience are shownin the figure below.
• CHECKOUT OF MAJOR SYSTEMS
- OPERATING ENVELOPE COVERED
- MODIFICATION TO HEATERS ON MODEL ACCESS SYSTEM
• INITIAL CALIBRATION
- MACH NUMBER UNIFORMITY
- TOTAL TEMPERATURE DISTRIBUTION
- TOTAL PRESSURE DISTRIBUTION
• INITIAL MODEL TESTS
- PATHFINDER I
- SPACE SHUTTLE
312
An a e r i a l  view o f  t he  Nat ional  Transonic F a c i l i t y  (NTF) s i t e ,  viewed from 
t h e  back side, i s  presented. The l a r g e  bu lk  storage n i t r o g e n  tank on the  l e f t  
and the  vent  stack a t  t he  r i g h t  o f  the tunnel  are used t o  support  t he  
cryogenic mode o f  operat ion.  The c o o l i n g  tower i n  the  foreground i s  used w i t h  
a water-cooled heat  exchanger i n s i d e  the  tunnel t o  support  t he  a i r  mode o f  
operat ion.  The h i g h  bay b u i l d i n g  i n  the  background houses model p repara t i on  
bays and shop area on the  f i r s t  f l o o r  and the c o n t r o l  room and tunnel t e s t  
s e c t i o n  entrance on the  second f l o o r .  The design performance c a p a b i l i t y  i s  a 
Mach number range o f  0.2 t o  1.2, pressure range o f  1 t o  8.8 atm, and 
temperature range o f  77 t o  339K. Th is  w i l l  produce a maximum Reynolds number 
o f  120 m i l l i o n  a t  a Mach number o f  1.0 based on a chord l e n g t h  o f  0.249 
meters. 
31 3 
LN2 PLANT 
The m a j o r i t y  o f  the  n i t rogen  i s  supp l ied  t o  the  o n s i t e  946 cub ic  meter 
s torage tank by p i p e l i n e  from a commercial a i r  separat ion p lan t .  
has been opera t iona l  s ince  January 1983 and supp l ied  most o f  the  l i q u i d  
n i t r o g e n  f o r  the  iJTF dur ing  checkout. The o n s i t e  storage tank can a l so  be 
f i l l e d  from mobi le  t ruck  u n i t s .  
Th i s  system 
ORIGINAL PAGE. 1s 
DE POOR QUALI'W 
31 4 
The sa fe  vent ing  of t he  gaseous n i t r o g e n  t o  the  atmosphere i s  c l o s e l y  
moni tored du r ing  a l l  c ryogenic  t e s t  runs. 
a i r  w i t h  the  gaseous n i t r o g e n  expe l led  from the  tunnel  i n  a vent  stack 37 
meters high. 
cond i t i ons  so t h a t  t h e  oxygen conten t  a t  t h e  s tack  e x i t  i s  a t  l e a s t  t e n  percent  
by volume. 
v i s i b l e  plume emi t ted  i s  dependent on the  atmospheric humid i ty  and wind 
cond i t i ons .  
A f an /e jec to r  system mixes ambient 
The m ix ing  r a t i o  i n  the  stack i s  a t  l e a s t  one t o  one under a l l  
The temperature i s  s t i l l  low a t  t he  e x i t  so t h a t  t he  s i z e  o f  t he  
31 5 
REYNOLDS NUMBER ENVELOPE
The maximum operating envelope showing Reynolds number as a function of
Mach number is presented and gives a fairly complete picture of the range over
which the NTF has been operated. The boundary lines and the lines of constant
total and dynamic pressure correspond to operation at minimum cryogenic
temperature. The maximum Mach number achieved thus far in the checkout and
calibration phase was 1.22. This Mach number was obtained in nitrogen at a
pressure of 1.2 atm and at a temperature of 180K. In air, the maximum Mach
number was 1.12 at a pressure of 1.4 and at a temperature of 320K.
q (atm)
120 x 10 e 3.3
100
80
R_.
6O
4O
2O
.8 1.0 1.2
M
316
SURVEY RAKE
The survey rake was utilized to determine total pressure and temperature
uniformity in the test section. The rake had an array of twenty total
pressures and ten total temperatures and was mounted in the NTF roll coupling
(internal to the model support arc sector) to provide roll capability of
+ 180 o . The front of the rake was located at the arc sector center of
_otation (station 13). The pressures were measured with the ESP system, and
the thermocouples utilized an onboard reference junction to enhance
measurement accuracy. The absolute level of the thermocouple measurements was
anchored by a platinum resistance thermometer mounted in the settling chamber
of the tunnel.
-->,.I.15 MI_--
20 TOTAL--_
PRESSURES _
10 TOTAL
TEMR
;- -4--
U i
"__IC=10_
A-.___AA -I"RAOIUS
CORNERS
BLOCKAGE AREA RATIO 0.5%
PLANFORM AREA RATIO 5.6%
317
TOTAL PRESSURE DISTRIBUTION - 324K
Typical variations of total pressure across the test section normalized
by the reference total pressure for air at 324K are presented. The reference
line faired through the data is for a condition of zero gradient. It will be
noted from an inspection of the data that a pressure gradient is not
detectable in either the horizontal or vertical direction. Data taken at other
radial cuts across the test section indicate the same result.
ROTARY RAKE PRESSURES
PT
PREF
1.01
1.00
.99
1.01
1.00
.99
PREF : 1,21 BARS
TT : 324K
HORIZONTAL
SYM MACH ROLL
0 0.999 0.0
D 0.994 18_.0
0 0.996 90.0
0.993 270.0
L
_ _/'-VERTICAL
1,1,1,1,1,1,1,1,1,1,1,1
-0.4 -0.2 0.0 0.2 0.4 0.6
r
1/2H
318
TOTAL PRESSURE DISTRIBUTION - 122K
Typical variations of total pressure across the test section normalized
by the reference total pressure for cryogenic operation at 122K are
presented. The reference line faired through the data is for a condition of
zero gradient. It will be noted from an inspection of the data that a
pressure gradient is not detectable in either the horizontal or vertical
direction. Data taken at other radial cuts across the test section indicate
the same result.
ROTARY RAKE PRESSURES
PT
PREF
1.01
1.00
.99
S
1.01
1.00
.99 --
-- _/--VERTICA. L
_,l,l,Iil,l_Iiltll
-0.6 -0.4 -0.2 0.0 0.2
r
112 H
PREF : 2,11 BARS
TT = 122K
SYM MACH ROLL
O I .024 0.0
I"1 1.036 t80.0
I .030 90,0
A I,010 270.0
319
TOTAL TEMPERATURE DISTRIBUTION - 324K
Typical variations in total temperature across the test section
normalized by the reference total temperature are presentea. The data were
taken in air at 324 Kelvin using the cooling coil. The reference lines
represent zero gradient. The distribution of data about this line is within
plus or minus 1 Kelvin.
ROTARY RAKE TEMPERATURES
T T
TREF
/HORIZONTAL TREF 32q,3 K
1.O3 _ jQ PT = 1,21 BARS
1.02 _" un,, n n_ o o o o _ vo u u a o u _ SYM MACH ROLL
1"O 1 I---- 1K ._T o 0.999 0.0
>
D 0,994 180.0
O 0.996 90.0
0,993 270.0
1.03 --
1.02 r O"o .C,
1.01 -- ^_^ ± _Q" _ L'T"1K
,I,I,I,I,I,I,I,I,I,I,I
-0.4 -0.2 0.0 0.2 0.4 0.6
r
112 H
320
TOTAL TEMPERATURE DISTRIBUTION - 122K
Typical variations in total temperature across the test section
normalized by the reference total temperature are presented. The data were
taken in the cryogenic mode using the liquid nitrogen injectors for cooling.
These data generally fall within a band of plus or uninus 0.5 Kelvin. It
should be noted that the total temperature ratio is slightly higher than
unity. This is due to an offset in the reference temperature junction which
was not corrected during the test. This is not important for this discussion
since the temperature uniformity across the test section was the primary item
of interest and is not affected.
ROTARY RAKE TEMPERATURES
TT
TREF
1.01
1.00
.99
HORIZONTAL
--- "-o? .... o_ .... ____ -____ -- 7-"-oo.... .... o
rl _
VERTICAL
_----_----_--- _ z_
•"- -" ..... _--z_- -A.... z_.....
- -0- __<>__<>__?o _<>__A_ ,,<>
r 1K
_ I, I, I, IJ I, I_ I, I,l,llli I
-0.6 -0.4 -0.2 0.0 0.2 0.4 0.6
r
1/2 H
16-31
TREF = 122,3 K
PT = 2.11 BARS
SYM MACH ROLL
O 1.024 0.0
D 1 .036 180.0
O I .030 90.0
A 1.0113 270.0
321
ORIGINAL PAGE IS 
OF POOR QUALW 
PATHFIWEK I 
The model represents  a gener ic  t r a n s p o r t  w i t h  s u p e r c r i t i c a l  a i r f o i l  and 
aspect r a t i o  o f  9.8. 
o f  t h e  model deformat ion system. 
Spots on the  l e f t  wing are  t a r g e t s  used i n  the  checkout 
322 
TUNNEL FLOW ANGLE
The tunnel flow angularity was investigated using the Pathfinder I model
to obtain an integrated value of the flow angle by testing the model upright
and inverted. The variation of normal force coefficient with model angle of
attack for the model upright and inverted is shown for Mach numbers of 0.60 to
0.82. The near perfect agreement between the upright and inverted runs at
both Mach numbers indicates that a correction for flow angle in the NTF at
these test conditions will not be required.
O N
.6
.5
.4
.3
.2
.1
0
-.1
MODEL
0 UPRIGHT
[] INVERTED
M=0.82
M-0.60
PT = 4.14 BARS
TT = 320 K
--DATA UNCORRECTED FOR FLOW ANGLE
--NO SIDE SLIP ANGLE
I I I
-1.0 -.5 0 .5
MODEL PITCH, DEGREES
I I I I
1.0 1.5 2.0 2.5
323
TEST ENVELOPE FOR MODEL INSTRUMENTATION CHECKOUT
For purposes of checking the effects of temperature on model
instrumentation performance, it was desirable to have a condition where the
flow would be sensitive to Reynolds number and the same test Reynolds number
could be obtained at ambient and cryogenic conditions. The Pathfinder I model
was used and tested over a Mach number range from 0.40 to 0.82. The test
condition of temperature and pressure is shown in the accompanying figure for
a Mach number of 0.70. A maximum chord Reynolds number of approximately 5
million was obtained in air at a temperature of 320K at 3 atm stagnation
pressure and the same Reynolds number was obtained at 177K and 1.33 atm
stagnation pressure. These results will be compared in subsequent figures.
Additionally, data will be presented to show the effects of Reynolds number
between 2.5 and 29.4 million by varying temperature from 320K to 117K at 1.33
atm stagnation pressure and then increasing pressure to 4 arm.
A Mach number of 0.70 was selected for analysis since any Reynolds number
sensitive separation present on the airfoil would not be dominated by shock
effects, and thus, would be less sensitive to any changes in model surface
conditions that might occur during the test due to particles in the stream.
The internal strain gage balance used to measure model forces and moments
was unheated. Temperature measurements were made at the bridges, and
corrections to zeroes and sensitivities were applied. For this balance, zero
shifts in axial force occurred during the test that were not experienced
either during calibration or cryogenic checkout prior to the test. Therefore,
drag data were not obtained. The zero shifts were repeatable in post test
checks and further evaluation is continuing. Angle of attack and pressure
instrumentation was maintained in a heated environment at all times.
M=0.70
PT
Psia
T,K 320 177 117
,oo / /
6O
40 / POINTS FOR WHICH DATA
ARE PRESENTED.
/ / /
20 tK)_/i /
o I I I I I I
0 10 20 30 40 50 60 70 80
Rc x 10 -6
324
EFFECT OF TEMPERATURE ON CN VERSUS e MEASUREMENT
The accompanying figure shows the variation of normal force with angle of
attack and indicates the same result is obtained at a constant Reynolds number
for a temperature of both 320K and 177K. The data for a Reynolds number of
2.5 million based on model chord show the familiar characteristics of the
supercritical wing tested at low Reynolds number. The nonlinearity of low
angles of attack has generally been attributed to flow separation on the lower
surface in the cusp region and, of course, trailing-edge separation is
apparent at the highest angles of attack.
PATHFINDER I ; M = 0.70
.8
CN
.6
.4
.2
T, K P, arm R x 10 -6
0 320 1.33 2.5
0 320 3.0 5.4
a 177 1.33 5.1
0
-.2 I I I I I
-2 0 2 4 6 8
, deg
325
EFFECT OF TEMPERATURE ON CM VERSUS _ MEASUREMENT
The accompanying figure illustrates the variation of pitching moment
coefficient with angle of attack for essentially the same Reynolds number
obtained at temperatures of 320K and 177K. The agreement is very good.
PATHFINDER I : M : 0.70
CM
.2
.1
0
-.1
T, K P, arm
0 320 1.33
0 320 3.0
I I I I
0 2 4 6
Rx 10 -6
2.5
5.4
5.1
I
8
o_., deg
326
EFFECT OF TEMPERATURE ON CM VERSUS CN MEASUREMENT
The accompanying figure illustrates the variation of pitching moment
coefficient with normal force coefficient for essentially the same Reynolds
number obtained at temperatures of 320K and 177K. The agreement between the
two conditions is very good.
PATHFINDER I ; M = 0.70
.2
I T, K P, a'tm R x 10O 32O 1.33 2.5
O
_ . 320 3.0 5.4
"1I _ _ 177 1.33 5.1
CM
0
--.1
--,2
-.2 0 .2 .4 .6 .8
CN
-6
I
1.0
327
EFFECT OF REYNOLDS NUMBER ON CN VERSUS
The accompanying figure illustrates the effect of Reynolds number on the
variation of normal force coefficient with angle of attack for the Pathfinder
I model. For this case, all of the effect of Reynolds number is obtained by 9
million. The small difference in angle of attack for a constant lift
(approximately 0.1 degrees) between 9.1 and 29.4 million appears to be the
result of a small shift in angle of attack zero. This is being evaluated.
1.0
PATHFINDER I : M = 0.70
.8
CN
.6
.4
.2
T, K P, arm R x 10 -6
O 320 1.33 2.5
117 1.33 9.1
A 117 1.33 29.4
0
-.2
-2 0
I I I I
2 4
_., deg
6 8
328
EFFECT OF REYNOLDS NUMBER ON CM VERSUS a
The accompanying figure illustrates the effect of Reynolds number on the
variation of pitching moment coefficient with angle of attack for the
Pathfinder I model. As in the case of normal force, essentially all of the
effect of Reynolds number is obtained by 9 million. These data also indicate
a probable shift in angle of attack between the 9.1 and 29.4 million Reynolds
number runs. At the highest angles of attack, a reduction in stability is
indicated for the 29.4 million Reynolds number due to aeroelastic effects on
the swept wing at 4 atm stagnation pressure.
CM
.2
.1
0
PATHFINDER I; M : 0.7
T, K P, al"m R x 10 -6
0 320 1.33 2.5
117 1.33 9.1
117 4.0 29.4
-.1
-.2
-2
i I I I I
0 2 4 6 8
o_.., deg
329
EFFECT OF REYNOLDS NUMBER ON CM VERSUS CN
The accompanying figure illustrates the effect of Reynolds number on the
variation of pitching moment coefficient with normal force coefficient for the
Pathfinder I model. The near perfect agreement between the data at 9.1 and
29.4 million Reynolds number indicates no change in load distribution between
the two cases. This supports the earlier statements that the differences in
normal force and moment with angle of attack are most likely due to a small
shifts in angle of attack zero between the two runs. The reduction in stability
due to aeroelastic effects is readily apparent for the case at 4 arm
stagnation pressure.
PATHFINDER 1; M: 0.70
CM
.2 -
.1
0
--,1
--.2
--.2
_JJ T, K P, a't'm R x 10 -60 320 1.33 2.5
117 1.33 9.1
I I I I
0 .2 .4 .6 .8
CN
I
1.0
330
SUMMARY
The NTF has been operated to design condition of 120 million Reynolds
number at a Mach number of 1.0.
All systems have been checked out except plenum isolation valves;
modifications are being made to heaters on the actuators.
Initial steady-state calibration indicates excellent steady flow
characteristics.
The first test of the Pathfinder I model indicated significant Reynolds
number effects.
Some effects of temperature on instrumentation were obtained. The
cause of these effects is being evaluated.
331


Preliminary calibration and test results from the National Transonic Facility

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