Welsh, Clement J

English

NASA Technical Reports Server

Copy 
-	 RML5OFO1 
NACA 
RESEARCH MEMORANDUM 
RESULTS OF FLIGHT TESTS TO DETERMINE THE

ZERO-LIFT DRAG CHARACTERISTICS OF A 60 0
 DELTA WING WITH

NACA 65-006 AIRFOIL SECTION AND VARIOUS DOUBLE -WEDGE

SECTIONS AT MACH NUMBERS FROM 0.7 TO 1.6

By Clement J
.
 Welsh 
Langley Aeronautical Laboratory

Langley Air Force Base, Va. 
EN(I E lIBRARY

CHANCE-VOUGHT AIRCIRAff

DALLAS, TEXAS 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
August 11, 1950
https://ntrs.nasa.gov/search.jsp?R=19930086188 2020-06-17T13:23:37+00:00Z
NACA RN L50FO1
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
RESULTS OF -FLIGHT TESTS TO DETERMINE THE 
ZERO-LIFT DRAG CHARACTERISTICS OF A 60° DELTA WING WITH 
NACA 67-006 AIRFOIL SECTION. AND VARIOUS DOUBLE-WEDGE

SECTIONS AT MACH NUMBERS FROM 0.7 TO 1.6, 
By Clement J. Welsh 
Results of an exploratory free-flight investigation at zero lift 
of several rocket-powered drag research models equipped with 600 swept-
back delta wings are presented for a Mach number range from about 0.70 
to 1.60. The airfoil sections tested includedthe NACA 65-006 and a 
series of double-wedge sections with various thicknesses and positions 
of maximum thickness. 
The results of tests showed that, of the double-wedge sections with 
6-percent thickness ratio, the two sections with positions ofmaximum 
thickness at 20 and 50 percent of the chord had drag coefficients approxi-
mately equal through the transonic and supersonic Mach number range and 
had similarly occurring drag rises. The section with position of maximum 
thickness at 80 percent chord had a drag rise occurring at approximately 
0.15 Mach number lower than the drag rise of the other two sections. At 
M = 1.0, this section had drag coefficients more than twice those of the 
other two sections; however, this difference decreased with increasing 
supersonic Mach numbers. The wing drag calculated by the linearized 
theory was in qualitative agreement with the test results in indicating 
the effects of varying the position of maximumthickness. The double-
wedge section of 3-percent thickness ratio with position of maximum thick-
ness at 50 percent •chord had fairly constant drag coefficients throughout 
the supersonic region, which ranged from about '50 to 80 percent of the 
drag coefficients for the similar section with twice the thickness ratio. 
The theoretical wing drag for this section was in very good agreement with 
the experimental value. The NACA 65-006 section had lower drag coeffi-
cients throughout the test region than any of the double-wedge sections 
of the same thickness ratio, although at the highest Mach numbers covered 
by these tests, the differences became very small. 
Ulm
.2	 NACARM L70FO1 
IR0JCTI0N 
As part of the National Advisory Committee for Aeronautics program 
to determine the drag characteristics at zero lift of various wings at 
supersonic, transonic, and high-subsonic speeds, tests of a series of 
600
 delta wings with varying airfoil sections have been made. These 
tests were conducted at the Pilotless Aircraft Research Station at 
Wallops Island, Va.,. with the wings being mounted on rocket-propelled 
test bodies.,..	 . 
The results are presented as curves of total. drag coefficient and 
wing drag coefficient against Mach number. Curves of theoretical wing 
drag coefficients are shown, for the double-wedge-section wings, for 
comparative purposes.
SYMBOLS 
H	 Reynolds number based on wing thean aerodynamic chord 
C	 total.; drag coefficient based on exposed wing area 
wing plus wing-body interference drag coefficient based on 
W	 exposed wing area 
t/c	 wing thickness ratio 
t	 maximum wing section thickness 
c	 wing chord measured parallel to center line of body 
M	 Mach number 
M	 mass of the test vehicle, propellant expended 
dV/dt . rate of change of velocity along flight path
	 . 
acceleration due to gravity, 32.1740 feet per second per second 
7	 .	 flight-path angle, measured from horizontal 
P	 mass density of air 
S	 exposed wing area 
V	 velocity along, flight path
NACARML5OFO1	 3 
MODELS 
The general arrangement of the dragresearch'xnodels used in the 
present investigation is shown in figure .
 1 and a typical plan-view photo- 
graph is shown in figure 2. The body of the models was cylindrical with 
a pointed ogival nose and was stabilized with four thin fins, located 
near the base. The wings investigated were of delta plan form, had 
600 sweepback of the leading edge, were of equal size relative to the 
body, were mounted on the body in the same location, and differed only 
in airfoil section. The variations of the airfoil sections of the five 
configurations investigated are indicated in the table shown in. figurel. 
Three configurations were double-wedge airfoil sections of 6-percent. 
thickness ratio but having the position of maximum thickness of the sec-
tion located at 20, 50, and 80 percent of the chord. A fourth configura-
tion was a double-wedge, 3-percent-thick section with maximum thickness 
at 50 percent chord. The last configuration had an NACA 65-006 airfoil 
section. Models without wings were flown to make possible the determina-
tion of the increment in drag produced by addition of the test wings. 
For convenience, the double-wedge sections with position of maximum thick-
ness at 50 percent chord will be referred to as symmetric sections in the 
rest of this paper. 
The bodies of the models were made of pine, and balsa wood, and the 
wings and fins were made of aluminum. The models were propelled as two-
stage-rockets. The first stage or booster employed a 5-inch,high-velocity 
aircraftrocket. The models comprised the second stage and were propelled 
by a 3.25-inch Mark VII aircraft rocket within the model. 
TESTS 
The models'were flown at the Pilotless Aircraft Research Station, 
Wallops Island, Va. The tests were performed by the 'usual method. The 
models were launched at an elevation angle of approximately 70° 'above 
the horizontal and drag measurements made during the coasting period of 
the model down through the Mach number range to subsonic speeds. From 
the summation of the forces acting upon the model along the direction 
of the flight path, the drag force may be found and equated to the 
standard formula for drag involving the drag coefficient, thus giving 
C	
m(dV 
 D	
pSV2
NACA RN L5001 
The range, velocity, and acceleration relative to a point on the ground 
near the launcher were measured from the ground by a continuous-wave 
Doppler type radar. The trajectory was measured with an SCR-584 radar 
theodolite., The trajectory measurements provided . the ,
 flight-path angle 7, 
the altitude, and the small corrections to the measure .d velocity and 
acceleration necessitated by the slight curvature of the portion of the 
trajectory during which the drag measurements were obtained. The varia-
tion of atmospheric pressure and temperature with altitude, which gives 
the air ' density p, the velocity of sound for determining the Mach number, 
and the viscosity of air for determining the Reynolds number, was measured 
by radiosonde at the time of the tests. 
The continuous-wave Doppler type radar furnished a time history of 
the radial distance to the model. Velocity and acceleration were 
obtained, respectively, as the first and second time derivatives of this 
time history. The method by which the two differentiations were obtained 
has been analytically developed to its present state of precision which 
is less than 0.7 foot per second velocity error and less than 3 feet per 
second per second acceleration error. 
The wing drag coefficients, including wing-body interference, were 
obtained as the difference between the drag coefficients of winged and 
wingless models. The tests were performed with the wings mounted on a 
readily constructed body which had drag coefficients that were well-
established from previous tests. The difference between the drag coef-
ficients of the winged and of the. wingless models being small relative 
to the drag of wingless model, particularly at subsonic speeds, causes 
low accuracy of the determined wing-plus-interference drag coefficients; 
however, the accuracy is sufficient for displaying the trends sought in 
this exploratory investigation. Because of the relatively low accuracy 
required in this exploratory investigation, repetitive tests were per-
formed in only a few cases, so that assurance is not had that, in the 
single tests; the results do not deviate from the correct values to an 
extent greater than the amount normally existing in repetitive tests of 
this type. From a large number of similar previous tests, the probable 
error in wing drag coefficients is estimated to be ±0.002 at. M = 0.80, 
±0.0013 at M =1.1, and 0.0035 and -0.0015 at M = 1.4. The probable 
error in Mach number is estimated to be ±0.01 at M = 0.8 and ±0.005 at 
M = 1.14. 
The average Reynolds number of the ten models tested,based on wing 
mean aerodyn9ic' chord of 15.25 inches, varied from 3.5 x io6 at M = 0.61 
to l4-.2 x 10° at M = 1. 75. A plot of Reynolds number against Mach 
number is shown in figure 3.
NACA RN L50FO1	 5 
RESULTS AND DISCUSSION 
The total drag coefficienté CDj plotted against Mach number M. 
for all the configurations investigated. including the basic wingless 
body are presented in figure 	 Two models were flown for each of the 
five configurations tested, however, data were obtained for only one 
model for each of the two symmetrical, .double-wedge section configura-
tions. The drag of the wingless body was subtracted from the total 
drag of each configuration leaving the wing dragplus wing-bodyinter-
ference drag for each. This wing drag coefficient is shown plotted 
against Mach number M in figures 5 and 6. 
Calculated. wing drag coefficients are also shown in figures 5 and 
6 for the double-wedge sections. The calculated values include a con-
stant viscous drag coefficient, estimated at 0.006, which has beèn 
added to the theoretical wave drag coefficients obtained.from reference 1. 
Of the double-wedge sections with 6-percent thickness ratio, the 
two sections with positions of maximum thickness at 20 and 50 percent 
of the chord had drag coefficients approximately equal through the 
transonic and supersonic Mach number range and had similarly occurring 
drag rises. The section with position of maximum thickness a.
 t 80 percent 
chord had a drag rise occurring at approximately 0.15 Mach number lower 
than the drag rise of the other two sections. At M = 1.0, this section 
had drag coefficients more than twice those of the other two sections; 
however, this difference decreased with increasing supersonic Mach 
numbers. The wing drag calculated by the linearized theory was in quali-
tative agreement with the test results in indicating the effects of 
varying the position of maximum thickness. 
The symmetrical, double-wedge section of 3-percent thickness ratio 
had fairly constant drag coefficients throughout the supersonic region, 
which ranged from about 50 to 80 percent of the drag coefficients for 
the symmetrical, double-wedge section of 6 percent thickness. The theo-
retical wing drag for this section was in very good agrement. with the 
experimental value. 
The NACA 65-006 section had lower drag coefficients throughout the 
test region than the symmetric double-wedge section of the same thick-
ness ratio. In the region of M = 0.975, the NACA 65-006 se tlon appears 
to show a favorable wing-body interference drag. 
- .
6	 .	 -.	 .	 NACA.RM L5OFO1 
CONCLUDING REMARKS 
Measurments of the effect of airfoil section on the wing plus wing-
body interference drag at zero lift of delta-plan-form wings having. 
600
 leading-edge sweepback and tested at high Reynolds numbers over the 
Mach number range from about 0.7 to 1.6 in free flight on cylindrical, 
fin-stabilized bodies with pointed nose lead to the followingconclusions. 
1. Of the double-wedge sections with 6-percent thickness ratio, the 
two sections with positions of Iriaxinlum thickness at 20 and 70 percent of 
the chord had drag coefficients approximately equal through the transonic 
and supersonic' Mach number range and had similarly occurring drag rises. 
The section with position of maximum thickness at 80 percent chord had a 
drag rise occurring at approximately 0.15 Mach number lower than the drag 
rise of the other two sections. At M = 1.0 1 this section had drag coef-
ficients more than twice those of the other two sections; however, this 
difference decreased with increasing supersonic Mach numbers. The wing 
drag calculated by the linearized theory was in qualitative agreement 
with the test results in indicating the effects of varying the position 
of maximum thickness. 
2. The symmetric, double-wedge section of 3-percent thickness ratio 
had fairly constant drag coefficient throughout the supersonic region 
which ranged from about 50 td 80 percent of the drag coefficients for 
the symmetric, double-wedge section of 6-percent thickness ratio. The 
theoretical wing drag for this section was in very good agreement with 
the experimental . value. 
3. The NACA 65-006 section had lower drag coefficients throughout 
the investigated transonic and supersonic regions than any of the double-
wedge sections of the same thickness ratio, although at the highest Mach 
numbers reached, the differences became very small. 
Langley Aeronautical Laboratory 
National Advisory Committee for Aeronautics 
Langley Air Force Base, Va. 
REFERENCE 
1. Puçkett, Allen E; and Stewart, H. J.: Aerodynamic Performance of Delta 
Wings at Supersonic Speeds. Jour. Aero. Sci., vol. 14, no. 10, 
Oct. 1947, pp. 567-578.
L=^_'_j
-1 
Cd 
4H H Q) 
'ti
Ui 
H •r-
H U) 
011) 
cO 
a) .,-4 
11)-f) 
a)0 
U) 
cH 
-i •r-4 
rd  
4-4 ^-q a) 0 r 
-p
Cd 
H r-1 
a).H 
U) 
Hc-i 0 (3 H •H 
(I) 
a)
Ea a) 
e' a. 4 rd 
I U) 
Ha) Ha) 
-4	 (	 1
7 
tn 
0 
0• 
/
NACA RM L50FO1
0 
11) 
c'J 
0 
0 
C)
() C.D to to iitti 
LL
Rq
Page intentionally left blank 
Page intentionally left blank
NACA RN L50FO1 
Figure 2.. Test vehicle showing plan view of delta wings investigated.
Page intentionally. left blank 
Page intentionally left blank
NACA BM L70FO1	 11 
I riu--------
R8 
4 
I-)
	 I	 I 
•	
.6	 .8	 /0	 /4	 16	 1.8 
Iv 
Figure 3.- Average variation of Reynolds number with Mach number for all 
models tested., based on mean aerodynamic chord of the wing, 15.25 inches. 
S
12	 NACA RM L50FO1 
.08 
ff
0 
.08
• ___A
MEN 
 _ 
NEON
 
12	 14	 16 
41 
0 
C12r .01
One model
16 
M 
.08 
C ,7- 
. 04
	 H 
Two mode/s 
t/c0.06 ZZ7 
I,) 
•	 .6	 .8	 /0	 1.2	 1.4	 1.6	 /8 
M 
Figure 14• Variation of total drag coefficient with Mach number. Wing of 
200 square inches; aspect ratio, 2.31; sweepback angle of leading 
edge, 60°.
NACA RM L50FO1
	
Mill
	
13 
CJJ,7-.04 
H
	
Omm 
•	 .6	 .9	 /6'
	
9 
41 
.08 
CDT
IVAcA 
<zz - Two mode/s 
65-006
16
	 9 
41 
.08 
cDT .01 
0 
•	
.6	 .6'	 /0	 /2	 14	 16	 1.8 
41 
Figure 4, Concluded.
NACA RM L50FO1 
f=0.06_/  
- -	
- --:- -	 •-- -
---- _____ _____ _____ 
•
1er,rnenf	 ,• - 
Thcy	 - 
/,606 a:___
==_ '
iT-- 
_ _- ------
___ 0u6
/
____ -
	
.8	 /0	 14	 16	 /8 
(a) Expeximenta1 and theoretical wing drag coefficients. 
08 
	
6'W.04
	
.1	 /	 ___-	
•.
a06 
r/J 
	
• .6	 .8	 /0	 12	 14	 16	 -.	 1.8 
41	 -	 AA 
(b) Summary plot of experimental wing drag coefficients. 
Figure 5.- Comparison of the wing drag coefficient of the double-wedge 
sections of 6-percent thickness ratio with varying positions of 
maximum thickness. The coefficients are based on wing area of 
200 square inches. 
114.
• .04 
0 
.04 
low
.04 
0
.6
NACA 1*4 L50FO1
	 15 
94	 I
Fx,oerimen/ 
Theory 
94
•i o6 	
-----L---	 --
/2L to	 te	 /1	 18 
41
(a) Experimental and theoretical wing drag coefficients. 
.08 
• __ __ __ __ __ __ JI t/=aQ6  
/lzzz-
_____
	
65-006 
//ç003 
14	 16 
(b) Summary plot of experimental wing drag coefficients. 
Figure 6.- Comparison of the wing drag coefficients of the NACA 65A006 and 
two symmetrical double-wedge sections with different thickness ratios. 
The coefficients are based on wing area of 200 square inches. 
NACA-Langley - 8-11-50 - 325


Results of flight tests to determine the zero-lift drag characteristics of a 60 degree delta wing with NACA 65-006 airfoil section and various double-wedge sections at Mach numbers from 0.7 to 1.6

Results of flight tests to determine the zero-lift drag characteristics of a 60 degree delta wing with NACA 65-006 airfoil section and various double-wedge sections at Mach numbers from 0.7 to 1.6

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server