NATIONAL AERONAUTICS AND SPACE ADMINISTRATIONArnold Engineering Development Center Arnold Air Force...
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P I I I i I' NATIONAL AERONAUTICS AND SPACE ADMINISTRATION -7 DMS-DR-2154 NASA-CR-134,437 HEAT TRANSFER TESTS OF A 0.0175-SCALE SPACE SHUTTLE ORBITER MODEL (29-0) TO DETERMINE THE EFFECT OF SURFACE TEMPERATURE ON BOUNDARY LAYER TRANSITION AT MACH 8.0 IN THE AEDC VKF, TUNAEL B (TEST OH4A) SPACE SHUTTLE AEROTHERMODYNAMIC DATA REPORT JOHNSON SPACE CENTER HOUSTON, TEXAS DNA LUNagernent services SPACE DIVISION 4, CHRYSLER CORPORATION https://ntrs.nasa.gov/search.jsp?R=19750007650 2020-06-19T01:28:55+00:00Z
Transcript of NATIONAL AERONAUTICS AND SPACE ADMINISTRATIONArnold Engineering Development Center Arnold Air Force...
P I I
I i I '
NATIONAL AERONAUTICS A N D SPACE ADMINISTRATION
-7
DMS-DR-2154 NASA-CR-134,437
HEAT TRANSFER TESTS OF A 0.0175-SCALE
SPACE SHUTTLE ORBITER MODEL (29-0) TO
DETERMINE THE EFFECT OF SURFACE TEMPERATURE ON
BOUNDARY LAYER TRANSITION A T MACH 8.0 I N THE
AEDC VKF, TUNAEL B (TEST OH4A)
SPACE SHUTTLE
AEROTHERMODYNAMIC DATA REPORT
JOHNSON SPACE CENTER
HOUSTON, TEXAS
D N A LUNagernent services
SPACE DIVISION 4, CHRYSLER CORPORATION
https://ntrs.nasa.gov/search.jsp?R=19750007650 2020-06-19T01:28:55+00:00Z
January, 1975
DMS-DR-2154 NASA-CR-134,437
HEAT TRANSFER TESTS OF A 0.0175-SCALE
SPACE SHUTTLE ORBITER MODEL (29-0) TO
DETERMINE THE EFFECT OF SURFACE TEMPERATURE ON
BOUNDARY LAYER TRANSITION AT MACH 8.0 I N THE
AEDC VKF TUNNEL B (TEST OH4A)
J. W . Foust S h u t t l e Aerosci ences
Rockwell I n t e r n a t i o n a l Space D i v i s i o n
Prepared under NASA Contract Plumber NAS9-13247
Data Management Services Chrys le r Corporat ion Space D i v i s i o n
New Orleans, La. 70189
for
Engineer ing Analysis D i v i s i o n
Johnson Space Center Nat iona l Aeronaut ics and Space Admin is t ra t ion
Houston, Texas
WIND TUNNEL TEST SPECIFICS:
Test Number: AEDC VKF Tunnel B Test VA352 NASA Series Number: OH4A Model Number: 29-0 Test Dates: Part 1 , 12 November 1973; Part 2 , 5 December 1973 3ccupancy Hours: 20
FACILITY COORDINATOR:
Mr. Larry Trimmer ARO, Inc. von Karman Gas Dynamics Faci l i ty Arnold Engineering Development Center Arnold Air Force Station, Tennessee
Phone: (615) 455-2611, E x t . 7377
37389
PROJECT ENGINEERS:
M. Quan (Part 1) W . R . Martindale J. W . Foust (Part 2 ) ARO, Inc. 12214 Lakewood Blvd. von Kaman Gas Dynamics Facil i ty Downey, California 90241 Arnold Engineering Devel. Center Mail Code: AC07 Arnold Air Force Station
Tennessee 37389 Phone: (213) 922-4826
Phone: (615) 455-2611 ext. 687
DATA MANAGEMENT SERVICES:
Prepared by
Reviewed’by: G. G . McDonald, J . L . Glynn
Liaison--D. A. Sarver, V . W. Sparks Operations--Maurice Moser, Jr.
Chrysler Corporation Space Division assumes no responsibil i ty f o r the data presented other than display character is t ics .
HEAT TRANSFER TESTS OF A 0.01 75-SCALE
SPACE SHUTTLE ORBITER MODEL (29-0) TO
DETERMINE THE EFFECT OF SURFACE TEMPERATURE ON
BOUNDARY LAYER TRANSITION AT MACH 8.0 IN THE
AEDC VKF TUNNEL B (TEST O H ~ A )
BY
J . W. Foust, Rockwell International Space Division
ABSTRACT
Test OH4A was conducted i n the Arnold Engineering Development Center
( A E D C ) von Karman Gas Dynamics Facility (VKF) Tunnel B.
heat t ransfer measurements on a 0.0175-scale model of the Space Shuttle
Orbiter w i t h the purpose o f determining the e f fec t o f wall temperature on
the point of boundary layer transit ion.
8 over a range o f Reynolds numbers from 1 .O t o 3.7 x 10 per foot.
I t consisted of
The t e s t was conducted a t Mach 6
Two regions of i n i t i a l model wall temperature were investigated.
Eighteen good runs were obtained a t each region i n support of the test
objectives.
scription o f the test procedure.
i n i t i a l model wall temperature did not obviously a f f ec t the point o f
boundary 1 ayer transit ion.
Tabulated data are presented i n this report along w i t h a de-
Preliminary data indicated that the
i i i
(THIS PAGE INTENTIONALLY LEFT BLANK)
TABLE OF CONTENTS
ABSTRACT
I N D E X OF MODEL FIGURES
I NTRODUCTI ON
NOMENCLATURE
REMARKS
CONFIGURATIONS I N V E S T 1 GATED
I N S T RL'MEN TA T I ON
F A C I L I T Y DESCRIPTION
T E S T PROCEDURES
DATA REDUCTION
DISCUSSION OF RESULTS
REFERENCES
TABLES
I. TEST CONDITIONS
11.
I I I.
DATA SET/RUN NUMBER COLLATION SUMMARY
MODEL DIMENSIONAL DATA
I V . THERMOCOUPLE GAUGE COORDINATES
V. LOCAL MODEL SURFACE ANGLES
FIGURES
MODEL
APPEND1 X
TABULATED SOURCE DATA
Page
iii
'2
3
4
8
10
11
12
13
15
18
19
20
21
24
30
31
32
35
Fi gure
1.
2.
3.
INDEX OF MODEL FIGURES
Title
Model 29-0 Nominal Dimensions
Chromel-Constantan Coaxi a1 Surface Thermocouple
Thermocouple Gauge Locations
2
Page
32
33
34
I ! c
c
INTRODUCTION
The surface of the Space Shuttle Orbiter i s subjected to changing
conditions, parti cul a r l y temperature, d u r i n g the cycle of launch, o r b i t ,
and return t o earth.
erature has on surface conditions.
layer t r a n s i t i o n is of in te res t since a change from a laminar t o turbulent
boundary 1 ayer represents higher convective heating rates.
I t is advantageous t o know the e f f ec t changing temp-
In par t icular , the point of boundary
This investigation ( t e s t OH4A) was undertaken t o determine the e f f ec t
o f model wall temperature on the point of boundary layer transit ion a t
i4ach 8.0.
temperature a t room temperature and approximately - 3 O O O F .
the investigation and t e s t procedures a r e presented i n this report.
Data were taken ( a t various Reynolds numbers) w i t h model i n i t i a l
The resul ts o f
3
Symbol
B
C
cP
F
h
H
hO
h
h / h o
h
h /h ,
h
h/ ho
K
M
P1 o t Symbol
CON FI G
GAGE
GROUP
HRE F-FR
H(TO)
N OMEN C L AT U RE
Definition
body
canopy or thermocoupl e gauge materi a1 specific heat (BTU/l b m - O R )
specific heat of tunnel airstream ( B T U / l b m - O R )
model configuration number
body f lap
model thermocouple gage number
run number
heat transfer coefficient ( B T U / f t -sec-OR)
Enthalpy (BTU/Jb)
2
theoret'cal s ta nation p o i n t heat transfer coefficient
heat t ransfer coefficient a t r = 1 (BTU/ft*-sec-"R)
( B T U / f t 3 -sec-OR 3
H ( T O ) / H R E F r a t i o of heat transfer coefficient t o theoretical
H(.9TO)
H ( .9TO)/ HRE F
stagnation p o i n t heat t ransfer coefficient
heat transfer coefficient a t r = .9 (BTU/ft2-sec-"R)
r a t i o of heat transfer c o e f f i d e n t t o theoretical stagnation point heat t ransfer coefficient
heat transfer coefficient w i t h r as a unction o f the local flow deflection angle (BTU/ft -sec-"R) i! H(TAW)
H(TAW)/ HREF
r a t i o of heat transfer coefficient t o theoretical stagnation p o i n t heat t ransfer coefficient
thermocoupl e gauge materi a1 therma 1 conducti v i t y (BTU/ft-sec-OR)
OMS pod
4
d
Sym bo 1 M
'm
P
Pm
r
R
Re
'm
S T
t
T
Tm
Tw
m V U
V
P i 0 t Symbol MACH NO
M U - I N F
P - I N F
P O
Q - I N F
Q DOT
RE/FT
RH& I N F
STFR
T- I N F
TO
Tw
V.- I N F
NOMENCLATURE (Continued) - Defi nXlori
frees tream Mach number
freestream viscosity (Sb-sec/ftz)
pressure ( p s i a )
freestream s t a t i c pressure (ps ia )
freestream stagnation pressure (ps ia )
freestream dynamic pressure (ps ia )
heat transfer rate (BTU/ft2-sec)
recovery factor
r a d i u s of a scaled one-foot sphere ( f t )
gas cons tan t ( f t - lb / s l ug-OR)
freestream u n i t Reynolds number ( l / f t ) freestream density ( s lugs / f t 3 )
Stanton number
time (sec)
temperature or temperature difference ( O R )
freestream s t a t i c temperature (OR)
freestream stagnation temperature (OR)
model surface temperature (OR)
freestream velocity ( f t / s ec )
velocity i n the x direction a t the edge o f the boundary layer vertical t a i l
5
NOMENCLATURE (Continued) Plot . --
Symbol Symbol Definition W wing
X
X/L X/L axi a1 1 ocation of thermocouple gauge
2Y/b 2Y /B span wise location of thermocouple gauge
distance along the curved surface of a sphere ( f t )
A
!J
i ncremen t
viscosity of a i r (1 b-sec/ft ) 2
IT cons t a n t
P density (slugs/ft3) - . -
T dummy time variable of integration
Anql es
" ALPHA- madel angle of attack MODEL
"S ALPHA- st rut pitch angle SECTOR
ALPHA- s t i n g o f fse t angle PREBEND
AW
i
local flow deflection angTe
local model surface angTe
PH I
ROLL- model rol l angle MODEL
thermocouple gauge rol l angle location measured clockwise from the bottom centerline looking forward
PE AW mode 1 yaw an g 1 e
Subscripts and Superscripts
adiabatic wall conditions
sumnation index
6
c
W
NOMENCLATURE (Concluded) P l o t
Symbol Symbol D e f i n i t i o n i n i n i t i a l conditions
0 stagnation condit ions
W w a l l conditions
m frees tream condi t i ons
I condit ions behind a normal shock
7
Test
of runs w
a t attemp
REMARKS
OH4A was conducted in two parts. Part 1 consisted of a se r ies
t h t h e i n i t i a l model wall temperature near room temperature plus
t o complete the same ser ies of runs w i t h an i n i t i a l model wall
temperature of -300°F. Par t 1 was terminated d u r i n g the cold wall runs,
primarily because of a discontinuity a t the model nose/body j o i n t due to
thermal stresses and/or uneven thermal expansions. This discontinuity
caused inconsistent perturbations in the heating ra te data and also
affected the boundary layer transit ion point.
with the cold wall runs were f ros t formation on the model and microswitch
freezing i n the t e s t section t a n k while the model was cooled.
consisted o f the series of cold wall runs completed a f t e r changes were
made t o a l leviate the problems mentioned above.
t e s t r u n schedule are presented i n Tables I and 11,respectively.
Other problems associated
Par t 2
The t e s t conditions and
All problems encountered d u r i n g the cold wall portion o f Part 1 were
successfully resolved for P a r t 2 .
sides and lower surface to eliminate the discontinuity.
f ros t was prevented by the following three changes:
The nose/body j o i n t was welded onto the
The formation of
( 1 ) The model was injected into the airstream i n an upright position
instead o f inverted.
The liquid nitrogen manifold, used for cooling, was placed d i r -
ectly under the model a t the same angle,allowing the model t o
be bathed in nitrogen until injection i n t o the airstream.
( 2 )
a
REMARKS (Concluded)
( 3 ) During Reynolds number changes, the model was sprayed with
gaseous nitrogen a t zero degrees Fahrenheit.
cleaned the model of condensation, t h u s eliminating the need t o
open the tes t section tank, and also periodically warmed the
tank t o prevent microswi tch freezing.
This technique
c
9
CONFIGURATIONS INVESTIGATED
Model 29-0 i s a 0.0175-scale r e p l i c a of the Space S h u t t l e Orbiter de-
s igned t o t h e -139 lines. Dimensional d a t a f o r the model a r e shown i n
Table I11 and nominal dimensions a r e shown i n F igure 1. The model was
c o n s t r u c t e d of 15-5 PH s t a i n l e s s steel and s t i n g mounted through i t s base.
I t i s a thick skin h e a t t r a n s f e r model des igned t o accommodate pressure
t aps 1/4-inch diameter c a l o r i m e t e r gauges and 1/8-inch d iameter coax ia l
gauges. The rudder and elevon were n o t s imula t ed f o r th is model.
Nomenclature f o r model 29-0 i s a s fo l lows :
Fuselage body (Confi gura-tion 3) '17
c7 Canopy (Conf igura t ion 3)
F5 Body f l a p (Conf igura t ion 3)
M4 OMS pods (Conf igura t ion 3)
V e r t i c a l t a i l (Conf igu ra t ion 3) "7
"I03 Wing (Conf igura t ion 3)
10
i
4
!
INSTRUMENTATION
For t e s t OH4A, the model instrumentation consisted of seventy-five
AEDC 1/8-inch coaxial thermocouple gauges. The coaxial gauge is a surface
chrome1 -constantan thermocouple which relates surface temperature t o the
incident heat flux by homogeneous, one-dimensional , semi-infinite s l a b
theory. The coaxial thermocouple i s shown i n Figure 2.
The coaxial gauge locations a re shown in Figure 3 and are tabulated
All gauges are on the fuselage lower centerline, fuselage in Table IV.
lower surface (left-hand s ide ) , or l e f t wing lower surface.
A separate calibration was required for the thermocouple gauges a t
cold wall temperatures (-300OF).
c
11
FACILITY DESCRI PTI ON
The Arnold Engineering Development Center (AEDC) von Karman Gas Dy-
namics Facili ty (VKF) Tunnel 3 i s a continuous-hypersonic-closed c i r cu i t
wind tunnel w i t h a i r as the t e s t medium. I t has a 50-inch diameter test
section and i s capable o f nominal Mach numbers of 6 and 8.
injected into the t e s t section from a tank d i rec t ly below.
character is t ics o f Tunnel B can be found i n Reference 1.
Models are
Operating
12
I
TEST PROCEDURES
The ser ies of runs w i t h the i n i t i a l model wall temperature near room
temperature were completed f irst and followed, a f t e r a delay, by the series
of cold wall runs.
inverted i n t o the airstream for about three seconds to record data. I t
was inverted because of the blockage of the u p r i g h t ins ta l la t ion a t low
Reynolds numbers. The room temperature runs were completed without i n -
For the room temperature runs, the model was injected
cident and immediately followed by an attempt t o complete the cold wall runs
(see Remarks sect ion) .
the model, was located on the siaewali of the tunnel t e s t section tank,
The l i q u i d nitrogen manifold, used for cooling
which required t h a t the model be rolled w i t h the bottom surface facing
the manifold fo r cooling and then returned t o the inverted position for
injection.
on the model due t o the time delay.
This procedure contributed heavily t o the formation of frost
After the problems t-espunsible~fm tkextest de?ay were resolved, the
cold wall runs were completed. A new approach was used which allowed the
model t o be positioned upright ( a t 30" alpha) i n the test section tank
d i rec t ly over the l i q u i d nitrogen manifold.
injected into the airstream, and no blockage problems were encountered.
A cycling technique was devised dur ing the t e s t to decrease the cooling
After cooling, the model was
time of the model.
the l i q u i d nitrogen s tar ted to solidify.
First, the tes t section tank was evacuated u n t i i
Then the tank pressure was i n -
owly t o near atmospheric. T h i s cycle was repeated u n t i l the
temperature reached -3OO0F, a t which time the tank was evacuated
creased s
model wal
13
TEST PROCEDURES (Concl uded)
again t o ob ta in the p res su re d i f f e r e n t i a l needed t o open the tes t s e c t i o n
s a f e t y doors and i n j e c t the model (tunnel pressure 0.1 p s i a and tank
pressure 0.15 ps i a during i n j e c t i o n ) .
Normally two cyc les were requ i r ed be fo re the model s t a r t e d coo l ing ,
poss ib ly due t o gas purging from the l i q u i d n i t rogen system.
technique used was t o blow gaseous n i t rogen from t h e tank door onto the
model dur ing Reynolds number changes ( tunnel t o t a l pressure and t o t a l
t empera ture changes). The gaseous n i t rogen was a t approximately ze ro
degrees Fahrenhei t and served two purposes: ( 1 )
any condensation and ( 2 ) i t p e r i o d i c a l l y warmed the test sec t ion t a n k ,
w h i c h prevented microswitch f r e e z i n g and e l imina ted any need t o open the
tank door. I t was found t h a t w i t h longer model c l e a n i n g time, less con-
densa t ion developed on the model dur ing runs.
Another
i t c leaned the model o f
14
. 0
I c
DATA REDUCTION
Surface temperatures measured by the coaxial thermocouple gauge are
related t o the incident heat f l u x by homogeneous, one-dimensional , serni-
i n f in i t e slab theory.
( Reference 2 ) :
The heat flux i s expressed by the follawing
Where: T = TW - Twin
Values of T were calculated by converting gauge outputs i n mil l ivol ts
t o temperatures using curve f i t s of temperature versus mi l l ivo l t tables
f o r chrome1 -constantan thermocuupl es.
recorded on magnetic tape a t a rate of twenty times per second.
The thermocouple gauge outputs were
Reference (3 ) provides a concise numerical technique for o b t a i n i n g
heat f l u x f rom a surface thermometer o u t p u t . When applied t o Equation 1 ,
the solution is:
15
DATA REDUCTION (Continued)
Values o f 4(t) were calculated using Equation 2 and a time increment
of 0.05 second averaged over a period of one second.
Heat transfer coefficient, h , was calculated by:
h = 40 TAW-TW (3 )
Where: TAW = rTo
The recovery factor, r , was a function o f the local flow deflection
angle a t each gauge location.
where r = 0.9, the recovery factor i s defined by:
With the exception of gauge number 13,
r = TAW = 0.867 + 0.133 sin 1.55, TO
Where : 6 = a i o = local flow deflection angle.
a = model angle o f attack
(4)
o = local model surface angle
Values of o for each gauge are given i n Table V.
Heat transfer coefficient ra t ios , h/ho, were calculated using:
Where:
Wi th : -6 1 / 2 0.0232 x 10 T u =
P
. (1 + 220/T)
16
DATA REDUCTION (Concl uded)
. du = e) [ 2 RT ( 1 - Pm/Po)] 1 /2 dx
The Stanton number was calculated by:
ho ST = PwUwCp
The freestream conditions including Reynolds nurher were calculated
u s i n g equations shown i n Reference 4 w i t h measured tunnel stagnation
eratures, stagnation pressures, and s t a t i c 'pressures.
Raw data from the room temperature and cold wall ser ies of runs
reduced us ing two different coaxial thermocouple gauge calibrations.
was necessary due t o changes i n gauge properties, m a t the two d
ferent temperature regions.
(9)
temp-
were
T h i s
f -
A description of the tabulated data format can be found i n the t e s t
nomenclature.
1 7
DISCUSSION OF RESULTS
The t e s t OH4A tabulated data are presented i n Appendix A. Although
heat transfer coefficient d a t a are also presented for recovery factors of
1 .0 and 0.9, the primary in te res t here i s the heat t ransfer coefficient
data where the recovery factor varies w i t h the local flow deflection angle,
H(T*M)
Plotted data from t e s t OH4A are presented i n Reference ( 5 ) . The heat
t ransfer coefficient (H (T
plotted f o r various locations on the Orbiter fuselage and wing .
)/HREF) and temperature (Tw/To) ra t ios a re AW
Since the purpose of this report is t o present the t e s t data, no
conclusions can be stated regarding i t s content. However, i t can be s t a t e d
tha t the data obtained adequately met the t e s t requirements. The pre-
.
liminary data, reviewed a t the t e s t s i te , indicated no obvious e f f ec t on
the point of boundary layer transit ion due t o model wall temperature.
. i
18
REFERENCES
1. "Test F a c i l i t i e s Handbook," Arnold Engineering Development Center,
Ninth Edition, July 1971.
2. Carslaw, H. S. and Jaeger, J . C . , Conduction o f Heat i n Solids,
Second Edition, Oxford, Clarendon Press, 1959.
3. Cook, W . J . and Felderman, E . J . , "Reduction of Data from Thin -F i lm
Heat-Transfer Gages: A Concise Numerical Technique," AIM Journal,
Volume 4, No. 3, March 1966.
4. Ames Research S taf f , "Equations, Tables, and Charts f o r Compressible
Flow," National Advisory Committee f o r Aeronautics, Report 1135,
- %\ 1953.
b 5. W. R. Martindale and L . L. Trimmer, "Test Results From the NASA/
Rockwell International Space Shut t le Test (OH44) Conducted i n the ,:
J- AEDC-VKF Tunnel B," AEDC-DR-74-39, May 9, 1974.
19
TABLE 1. TEST CONDITIONS
Macn No. Total Pressure Total Temperature Reynolds No. M PO (psia) To ( O F ) Re/ ft
860
675
5k5
425
325
210
880
870
850
840
835
810
3.7 x 106
3.0 x lo6 2.5 x lo6 2.0 x 106
1.0 x 106
1.5 x 106
20
W t-
1
I t 1
t
i
21
T E S T RUN NUMBERS I a r >
22
.
1
TABLE 111.
MODEL COMPONENT
MODEL DIMENSIONAL DATA
BODY - &.-, - ,
D IMEN SlON 5 :
Length In.
Max Width In.
Max Depth - In.
Fineness Ratio
. Area Ft2
Max. Cross-Sectional
Planform
Wetted
Base
FULL SCALE MODEL SCALE
- J a s U L - 22.58025
267.6 4.6830
244.5 4.27875
1Rh.67 0 : 11842
24
c
TABLE 111. MODEL DIMENSIONAL DATA (COffll'D)
MODEL COMPONENT : CANOPY - C7 I
G E N E R A L DESCRIPTION C0-r ation 3 ~e r Rockwell Lines VL70-0oO139. I
DRAWING NUMBER ' m70-000139
DIMENSIONS : FULL SCALE MODEL SCALE
Length (q,=433 to &=578) - In. P . 1 1 5 , 2.538
Mox Width
Max Depth
Fineness Ratio
. Area
Max. Cross-Sectional
Planform
Wetted
Base
25
TABLE 111. MODEL DIMENSIONAL DATA (CONT'D)
MODEL COMPONENT : BODY FLAP - F r
GENERAL DESCRIPTION . 3 Confimration D er Rockwell Lines VL70-0001J9
~
0.0175
DRAWING NUMBER VL70-OOO139 *
DIMENSIONS :
Length - In.
Max Width - In. Max Depth
Fineness Rotio
Area Ft2
Max. Cross-Sectional
Planform
Wetted
Base
FULL SCALE MODEL SCALE
84.70 - 1.48225 A
38.0958
.
I
1
26
.
TABLE 111. MODEL DIMENSIONAL DATA (CORT'D)
CkWll Lines VL?O-O001~~ GENERAL DESCRIPTION * C0nfi.m r a w 7 Der Fo
ds 10-t f-
-~
MODEX SCALE: 0.0175
DRAWING NUMSER ' VL70-OOO139
DIMENSIONS : FULL SCALE MODEL SCALE
Max Width In. -- 111.0 1.493- Max Depth In.
Fineness Ratio
. Area
I
Max. Cross- Secti onal
P Ian form
Wetted
v Base
Cente r l ine of OMS Fbd:
WP - 463.9 114~s; WP 4i>o + 63.9 = 463.9 BP - 80 .0IWS Length 1214.0 to 1560.0 = 346.0 INTS
NOTE: $ i d e n t i c a l t o M of 2A conf igu ra t ion , except 3 i n t e r s e c t i o n t o body.
27
\
TABLE 111 DIMEXISIOEAL DATA ( CONT'D)
I RODEL COMPONENT: VERTICAL - V, I
GENERAL DESCRIPTION: C e n t e r l i n e b-er t ica l t a i l , doublewedge airfoil with
rounded leadina edge.
vr.7Q&QQQp& -95 . DRAWING NUMBER:
DIMENS IONS-: FULL -SCALE M O D E L SCALE
TOTAL DATA
Area (Theo) Ft2 P 1 an form
Span (Theo) In Aspect Ratio Rate o f Taper Taper Rat i o Sweep Back Angles, degrees
Leading Edge Tra i 1 ing Edge 0.25 Element Line
Chords: Root (Theo) WP ,
Tip (Theo) WP MAC Fus. Sta. o f .25 MAC W, P, o f .25 MAC
A i r f o i 1 Sec t i on Leading Wedge Angle Deg Trai l ing Wedge Angle Deg Leading Edge Radius
B. L. O f a 2 5 MAC .
Void Area B 1 anketed Area
10.000 + .2.0 11.17 n.nn
. .. ,
TABLE I11 DIMENSIONAL DATA - Concluded.
31 FENS IONS :
2 Area :??eo,) F t TCTAl 9ATA
? l a n f o m Span (The0 In, Asgec2 S a t i o 2ate o f Taper Taper Rat io Dihedra l Png?e, deg rees I n c Ydence Ar,gle , degrees Aerudynami c Twi st, degrees Sweep Sack Ang7 est degrees
Leading Edge T r a i 1 ing Edge 0.25 Element L i ne
R o o t (Theo B.P.O.O. TiD, (Theo 1 6.P. MAC Fus. S t a . o f .25 MAC W.P. o f .25 M A C B.L. o f .25 MAC
chords:
EXPOSED DATA Area tTheo) Ft2 Span, (Theo) In. BP108 Aspect Ratio Tape r Ratio chords
R o o t BP108 T f p 1.00 b
2 MAC Fus. Sta . o f .25 MAC W.P. o f .25 MAC 6.L. o f .25 MAC
FULL-SCALE MODEL SCALE
Airfoil S e c t i o n (Ftockwell Mod NASA) xxxx-64
Root b =
T j p b = P
Data for ( 7 ) of (2) Sides Leading Edge Cuff P ianform Area Ft Leading Edge I n t e r s e c t s Fus M. L. 8 S t a Leading Edge Intersects Wing @ Sta
7
29
0.10 0.10
0.12 0.12
3auge No.
I x / c 1 x Y/b/2 Y
FUSELAGE 4
I 52 ~
53 1 54 I 55 I 56 I 57 \ 58 59 60 1 61 , 62 I
63 64 65 66 67 68 1 69 70 71 72 73 74 I
75 , 76 77 I 78 79 8 0 1 81 82 83 a i 85 I
1.1
6 1
2 3 4 1
7 8 13 11 12 , 1 3 26 1'7 19 20 32 23 24 25 26 29 30 31 32
34 3s '
37 38 39 40 41 47 WI 45 ' 46 47 48 49 , 50 ;
51
33 ,
2.049 i
Y 3 278
1
~
Y 4.098
'i 4.918
1 'i'
6.147
P 6.147
f 6.967
1 4 098 7.786
1
0 0
.082 9.032 ' 0.25
.302 11.290 1 .020 -452 .045 -903 -012 .282
-447 13.548 -591 15.806 .736 18.064 ,881 20.322 -050 13.548 .loo u.022 ,200 14.910 .300 15.806 .560 18.064 .700 19.307 .900 21.066 .176 15.806
0 0 0 0 0 0 - - - 0 875 0 0
1 1
Y 0.40
1
.060
.080
.loo
.150 ,150 .150 .150 .200 .200 .225
1.355 1.806 2.258 3.387 3.387 3.387 3.387 4.516 4.516 5.081
20° 30" 45.5"
Y 0.50
18.064 '' 21.076 y . 15.995 0.m- 16.625 1
19.145 I 18.064 1
20.405 I 21.035 Y 17.084 0.75 18.064 19.022
20.962 f
17.801 0.85 18,578
19.618 0.50
19.762 , 1
c
1
19.992 0.75
19.355
18.792 0.95
20.871 I v
.250 5.645
.275 6.210
.NO '6.774
.300 6.774
.300 1 6.774
.300 1 6.774 ,325 7.339 e350 7.898
484 .900 . loo .m 430
,600 .m .m . loo 0300 .500 .700 .m .loo -300 .5oo * 700 .loo .500 .900
0 0
1.312 - 400 - 45"
0 0 875 0 0 875 0 0
.875 0 0
0
O i
0 1
.400
.400 e450 .500 .500 550 .600 .6OO .650 .7m 750
. N O
.800
.850
.900
.900
9 032 9.032
10.161 11.290 11.290 12-419 13 548 13 548 14.677 15.806 16.935 18.061, 18.064
20.322 20.322
19.193 .875 0 0 ,
-875 '
-950 21.451 0 ' 1.000 I 22.580 , 0 , 1.OOO 22.580 1.750 I
I I
* See Figure 1 for def in i t ion of X, Y, & 9
1 . 30
GAUGE NO.
1 2 3 4 6 7 8 10 11 12 16 17 19 20 22 23 2.4 25 26 29 30 31 32 33 34 35 37 38 39 40 41 43 44 45 46 47 48
8 (DEGREES)
L
TABU V. LOCAL HODEL SURFACE ANGLES
8 (DEXXIZS)
52 . 25 20.5 13. 5 27.5 10.0 7025 5. 50 3-25 3.25 3.25 20 0 2.0 10 5 1.25 100 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 100 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 1.0 -0.6 -2.0 -2.0
1 I i
49 50 5 1 52 53 54 55 56 57 58 59 60 61 62 63 64 65 66 67 68 69 70 71 72 73 74 75 76 77 78 79 80 81 82 83 84 85
-3 9'5 -4.5 -4.5
5 035 2. 5 2.65 2.25 0
-0.75 . 9.3 6.7 5.65 4.65 2.85 1.4 -4.75 4.55 5.0 -5.3 4.95 4.8 5 04 3.5
-1.55 -5.1
5 *3 5.25 4.7 1.5 -4.8 5.6 5.35 4.6 1.85 5.8 4.4 -4.3
I
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31
DIKEINSIONS ARE;' NOMINAL V h L ' I S
REAR VIEW
Figure 1 . - Model 29-0 Nominal Dimensions.
-- 32
Note: Shown 10 x sca le
I .D. 9.D.
Figure 2. - Chrome1 - Constantan Coaxial Surface Thermocouple.
33
- I
\ i 34 P
APPENDIX
TABULATED SOURCE DATA
35
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r ' . 1 - I
j
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9 2 0 3
z * v z I
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0 2 - 0 a 0
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5 0 0 0 0 5 ) 0 0 0 1 D O
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r- 0 . !
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c o o 0 0 0
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m m m II m n m n n.0 o o o o o o o o o c I I I 1 1 1 1 1 1 1
.......... U W N N N ~ N N N ~ '
c c.c m c
n c \ N
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a -
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a .e ., x s z IP n r J\ 1 J C 9 c oc (D o c o a r. a o o o c c o a o c c c 0 0 LL 1 1 - x o m u m c a o ~ ~ n m r - ~ ~ ~ c ~ - ~ ~ ~ 5 ? . - m m r - o - n ~ n ~ ~ c ' L . .................................. In -N
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0 .. ... OS9 * o m I-* , m e *
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O O O O O O C O O C O O O O O o t 0 O O o c c o o o c c ~ o o c c c o o c c c ~ ~ ~ N O O O C C O C Q C C ~ ~ ~ ~ ~ C C C ~ ~ ~ C C ~ ~ C ~ C C ~ ~ C ~ ~ ~ ~ ~ ~ ~ -. c-N .t Q r o r .n c 5 NW. c o a 5 o NIT. c c m o o m c otr or^ c c c o = m e o X 0 0 0 0 0 0 - ~ ~ - N N N N N ~ ~ ~ ~ ~ ~ ~ U e ~ ~ ~ Q ~ Q c ~ ~ ~ ~ ~ U 9 0 0 ........................................
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o o o o o d o o o o o d o o o o o d o o o o o d a O O U O O O O O O O O O O O O O O O O O O O O Q O o o o o o a o o o o ~ o o o m m n m m n ~ , n ~ , ~ n J U ~ * V ~ ~ Y , ~ Q Q ~ Q Q ~ - ~ ~ C C ~ ~ ~ ~ ~ O . ......................... V
4
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c o o o o o c o o o o o c c c c o d o c c c c c e 5 0 E C ~ 9 0 C C C e C C C C C S c o ~ C C C f C C 9 0 C C 3 c 0 5 C m O C C C O e 0 0 0 o o o o o m v 1 r-o - ur- o - N rt 9 s o - m a b o - m m - m o .........................
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IC u * e44 5 Jc c rc ...... .T??
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a w \ o - h e r c x c m . - c c c n ~ c e c c ~ I V L C = o m o = . u oov. o m c3.3 J a m c c ~ r n m X O O O O O O - - - - N N N N N ~ ~ ~ ~ ~ ~ U ~ U ~ ) ~ ) ~ ) Q Q Q ~ I C ~ ~ ~ ~ ~ ~ ~ O c3. ........................................ * .i
t n T N
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u Am - P NC run cr- u IC C Q c cr- - N S N ?I s -r N - ~ w < . u U P I ~ o 4 et n m - u c N a Q L u n - e n , c u m n u , ~ n . - - I r . c u u n u u c c a c-nucr.e< N -a# I? O U I N or- m z - u I- UQ Q (F (F u o .L w Q UI m u1 m rn M ui Q c m z c- o -IU m - U F N - ~ - O - C C a c a c o c O O ~ C O C O C O C O C C C O C C - - - - - ~ ........................................
c . . . . . . . . . . .
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0 0 d 0 0 0 1 0 d 0 0 0 0 0 rn o o a o o o o o o o o o ~ o \ n n n n m n a o o o g c = . o * N W N ? W N * Y * J t J t d l 01 .............
0 0 0 0 0 d 0 0 3 0 0 d 0 0 0 0 0 d 0 0 0 0 0 0 0 0 0 0 0 0 d 0 0 0 0 0 ~ 0 o I E c 3 o ao- mmul in .n JI m n vl n ~ m ~ n ~ a 0 ~ - - - h c n r x ~ n ~ .............. . i - . - 4 - .
s V I : - c J
W '2 C I O a -I- C' X
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w u o L e I . u o r m -I 4
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F n f l N m Y w L T ........
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~ O o C O C O C O C o C - c c c c c o c o c c c c c = c c c c c c e c c e c o c
no o ~n c o m OOUI o m c o u o c m o c ~ m ~ ~ ~ ~ ~ a c c a ~ o a m e m o o ....................
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