Document
NATIONAL ADVISORY COMMITTEE
F O R A E R
WASHINGTON
-
, NATIONAL ADVISOFX COIQCCITEE FOR AERONAUTICS RESEARCH M E " D u M ~ MINIMUM DRAG O F FOUR VERSIONS O F A SWEPT-WING FIGRTEB AIRPLANE OBTAINED FROM FLIGHT TESTS OF ROCKET-BOOSTED MID3323 AT MACH MIMBERS FROM 0.81 To 1.71 I By Earl C . Hastings, Jr.
SUMMARY Tests conducted with four specific versions' of a swept-wing fighter airplane indicate that a large reduction in external-drag coefficient was accomplished by redesigning the original configuration.
The forebody modifications, which consisted of a smaller canopy, slimmer nose, and sharper inlet lip, reduced the value of the external- drag coefficient from 0 . 0 4 4 to 0 . 0 4 2 a t a Mach number of 1.05 and fYom 0.042 to 0 . 0 4 0 at a Mach nuuiber of 1.28. Recontouring this modified fuselage by increasing the cross-sectional area ahead of and behind the wing to obtain a more efficient area distribution at a Mach number of 1 . 2 resulted in an additional drag reduction throughout the Mach number range Values of external-drag coefficient fromthese tests were of the tests.
The drag-rise constant at 0.035 between Mach numbers of 1.05 and 1.71.
Mach number for each configuration was 0.93.
INTRODUCTION An investigation has been conducted by the Langley Pilotless Aircraft Research Division to determine the m i n i m drag of four spe- cific versions of a swept-wing fighter airplane. The first phase of the investigation was to determine the drag difference between two config- urations with different canopy and nose shapes. A f'urther investigation was then made to determine the minimum drag of the configuration rede- signed with an area-rule application for a Mach number of 1 . 2 (refs. 1 and 2 ) .
, A l l of these tests were conducted at the Pilotless Aircraft Research .
Station at Wallops Island, Va. with rocket-boosted models of the various configurations .
SYMBOLS A cross-section area, sq in.
longitudinal-accelerometer reading normal-accelerometer reading “ n / g - C mean aerodynamic chord chord-force coefficient, positive in rearward direction, CC a - A l l - @ ; qs drag coefficient, Drag/@ CD normal-force coefficient, positive toward top of model, CN acceleration due to gravity, 32.2 ft/sec 2 l 3 flight-path angle, deg Y k ratio of specific heats 2 length, in.
M Mach number ratio of mass flow through duct to mass flow through a stream
m l m o
tube of area equal to inlet-capture area under free-stream conditions static pressure, lb/sq ft P .
dynamic pressure, lb/sq ft n ’ r radius, in.
.
S total wing area (excluding chord extensions), sq ft 1 1 t time, sec 1 7 V velocity, ft/sec W weight, lb X station measured from nose, in.
Subscripts: mode 1 base choking-cup base duct exit duct inlet (capture) free stream ex external int internal tot total MODELS AND TESTS Figure 1 presents a three-view drawing of the final configuration tested. Since the primary difference in the four configurations was in cross-sectional-area development, only the one three-view drawing is presented; however the normal-cross-sectional-area distributions of the four configurations are shown as figures 2 to 6. The dimensional and mass characteristics are presented in table I .
Configuration 1 was a model of the first design proposal of a full- Configuration 2 had a modified forebody - an attempt to scale airplane.
reduce the supersonic drag level. The modifications consisted of a smaller canopy, a slimmer and slightly longer forebody with a sharper and a sharper inlet lip. Both of these models had internal flow nose, and were not instrumented. Photographs of configurations 1 and 2 are shown in figures 7 and 8.
.( ,, I - ' . a -
* . ..
Reference 1 indicates that the zero-lift drag-rise increments asso- ciated with wings near the speed of sound can, in some cases, be reduced by changing the axial distribution of the fuselage cross-sectional area.
Configurations 3 and 4 of the tests reported herein were developed on this principle by increasing the f'uselage cross-sectional area of con- 2 in front of and behind the wing in an attempt to further figuration reduce the drag at supersonic speeds. This redistribution of cross- sectional area was made on the ducted model without appreciably changing 6 ) .
the maximum cross-sectional area (fig. A method used to determine cross-sectional-area distributions for supersonic Mach numbers is dis- cussed in reference 2.
Configuration 3 had no underslung scoop inlet or internal flow and Both models 3 and 4 contained internal telem- configuration 4 was ducted.
eter systems to obtain flight data. Photographs of these models are shown in figures 9 and 10.
A l l of the models with internal flow (configurations 1, 2, and 4) had very similar ducting and bases. A single duct ran through the fuse- lage from the underslung scoop inlet to a single exit at the base. In order to choke the duct at flight Mach numbers greater than 1.0, each of these models had a choking cup installed at the duct exit. A sketch of the choking cup and base is also shown in figure 1 . A slotted total- pressure tube was installed in the duct of configuration 4 ahead of this choking cup in order to obtain data necessary to compute internal drag.
The telemetered models (models 3 and 4) were instrumented to measure normal and longitudinal accelerations and free-stream total pressure.
Configuration 4 was also instrumented to obtain values of duct total pressure. These models had four static-pressure orifices located on the base which were manifolded together to give an average static-pressure reading over the base. These quantities were transmitted from the model in flight to a ground receiving station where they were recorded.
Configurations 1 and 2 were boosted to Mach numbers of about 1.3 by using single 6.23-inch Deacon rocket motors. The Mach number range for configurations 3 and 4 was increased to about 1.7 by adding an additional ABL Deacon rocket motor to the booster stage. A photograph of one of the model-booster combinations is shown in figure 11.
A rawinsonde released at the time of firing obtained measurements of free-stream temperature, static pressure, and winds aloft. The veloc- ity of the models and their positions in space were determined by a C W Doppler radar set and an NACA modified tracking radar unit, respectively.
I ANALYSIS OF DATA
' L The CW Doppler radar set w a s used t o determine t h e t o t a l drag of a l l of t h e models during t h e decelerating portion of t h e f l i g h t . The method of analysis consists of differentiating t h e measured velocity Vrth respect t o ti= after correcting f o r flight-path angle and winds C, i s reduced f r o m t h e following a l o f t . The t o t a l drag coefficient t o t equation: %tot Reference 3 discusses the operation of t h e CW Doppler radar set and the method of data analysis i n more d e t a i l .
I n addition t o t h i s method of determining drag, configurations 3 and 4 w e r e instrumented so that normal-force and chord-force coefficients CN and C, could be computed. A comparison between chord-force coeffi- c i e n t s determined from the rocket model of configuration 3 and unpub- l i s h e d wind-tunnel values of minim-drag coefficient obtained f o r the same configuration (obtained in t h e Langley 8-foot transonic tunnel) showed agreement within the accuracy of the rocket-model tests. There- fore, the chord-force coefficients o f t h e rocket models were assumed t o represent the minimum-drag coefficients of the configurations.
I n order t o calculate the external-drag coefficient - it w a s first necessary t o compute the
- CDbase) (CDeXt - % t o t - 'Dint
base and internal-drag coefficients. On the telemetered models the data w e r e measured. Since the necessary t o calculate 'Dint and %as, i n t e r n a l ducting and base arrangements of configurations 1 and 2 were so similar t o those of configuration 4, values of CD and C i n t %ase obtained f r o m t h a t test were assumed t o apply f o r configurations 1 and 2 also.
A s mentioned previously, each of the ducted models had a choking Therefore the t o t a l base drag cup i n s t a l l e d at the duct exit ( f i g . 1).
of these m o d e l s w a s the sum of the base drag of the choking cups and the I n the tests of configurations 3 base drag of the models themselves.
over the base of the models w a s measured.
and 4, the s t a t i c pressure Because of the limited number of telemeter channels available however, t h e base s t a t i c pressure of the choking cup w a s not measured i n t h i s t e s t . Reference 4 presents data from a test where the base s t a t i c pres- sure of a similar choking cup was obtained. These values of choking-cup base pressure coefficient were assumed to apply to configurations 1, 2, and 4 of this investigation because of the similarity of the choking cup to the ones used herein.
The total base-drag coefficient of configu-
-
ration 4 was then computed as
- ( p c - po) (Choking-cup base area)
-(pb - P o ) ( Model base area)
C ss qs
I+[
for this configuration Since configuration 3 had no choking cup, ' % a s e was determined from the first term in the above equation.
With the instrumented model of configuration 4 it was also possible to calculate the internal-drag coefficient by the method of reference 5.
This method consists essentially of determining the loss in total momen- The tum of air flowing through the duct between free stream and exit.
equation used for computing C,, is as follows: int This coefficient could only be determined in this test for Mach numbers greater than 1.0, since at lower h c h numbers the duct was unchoked and all of the data needed to satisfy the above equation could not be obtained.
QUALITY OF DATA The quality of the Mach number and drag data presented in this paper is best illustrated by a comparison of the two sources o f data Both telemeter and collected from the tests of configurations 3 and 4.
tracking radar values of Mach number and total drag coefficient were available from these tests. The differences in these quantities as obtained by the two methods are presented at several Mach numbers in the following table : .
Configuration 3 Configuration 4 M = 1.03 M = 1.57 M = 1.10 M = 1-78 .
A Mach number o .015 0.010 0.010 0.010
a c ! 0.0010 0.0015 o .0015 0.0015
Dtot d Although no comparative data are available for models 1 and 2 because both Mach number and total drag coefficient were determined only from the tracking radar, it is believed that the quality of these data is as g o d as that for configurations 3 and 4.
from the two sources When fairing curves through values of c%"i; of data from configurations 3 and 4, values obtained fromthe accelerom- eter data were weighted more heavily since they were believed to be the more reliable data f'romthese tests.
Because of the similarity of the internal ducting and bases of the models with internal flow, the values of C and CD obtained Dbase int from the test of instrumented configuration 4 were assumed to apply to configurations 1 and 2 also. Even if a fairly large percentage of error in these values did exist between the configurations at supersonic speeds, this difference would have a negligible effect on the overall external- drag coefficients since the magnitude of the errors would be quite small.
TEST CONDITIONS The conditions for the four-rocket-model drag tests in terms of Reynolds number, trim normal-force coefficient, and mass-flow ratio are presented in figures 12, 13, and 14, respectively.
Reynolds number values (based on the length of the mean aerodynamic chord) are plotted against Mach number for each configuration in fig- ure 12. Values for configurations 1 and 2 are consistently larger at comparable Mach numbers than those for configurations 3 and 4 . This is primarily due to the lower altitudes at which tests of configurations 1 and 2 were conducted.
Trim normal-force coefficient C , is presented for the instru- trim These models were flown mented configurations 3 and 4 in figure 13.
with center-of-gravity locations of 6 . 9 5 and 6 . 5 6 percent of the mean aerodynamic chord, respectively. Agreement between the two curves is The transonic trim change is small, amounting to about 0.050 good.
Since configurations 1 and 2 were both between M = 0.93 and 0.99.
tested with center-of-gravity locations of 7.0 percent of the mean aerodynamic chord and were qdte similar geometrically to configura-
tions 3 and 4, values of c for configurations 1 and 2 are believed
%rim to be essentially the same as those shown for 3 and 4 in figure 13.
The mass-flow ratio m/% of ducted configuration 4 is presented in figure 14. Above M = 1.01 the values shown were computed from e . . e e e e . . e . .
e . . e e . . e . e e e: e . : &CA RM ~ 5 6 ~ 2 5 a
e. e.. e e e. e. . e e.. e. e.. e *
-‘ .
measured data.
At Mach numbers less than ‘1:O:it was possible to esti- mate by assuming that the static pressure at the duct exit was
4%
the same as that measured on the base of the model. This estimted curve is also shown in figure 1 4 along with estimated values of m / m , for configurations 1 and 2 computed at M = 1.0.
RESULTS AND DISCUSSION Figure 15 presents the total drag coefficient for configurations 1, 2, 3 , and 4. These values include the base drag of each model and the Data presented for configurations 1 internal drag of the ducted models.
and 2 were obtained only fromthe CW Doppler radar unit, but values of from both telemeter data and tracking radalL are shown for con- % o t figurations 3 and 4. The agreement between the two sources of data for configurations 3 and 4 is considered very good throughout the supersonic Mach number range.
The internal-drag coefficient as determined from the test of con- These values are small with a figuration 4 is presented in figure 16.
maximum C of 0.0010 occurring at M = 1.71. Between Mach numbers Dint Of 0.81 and 1.29, cDint is assumed to be zero. Also shown in figure 16
are values of Chase for configurations 3 and 4. Since configuration 3
was not ducted, the base geometry of this configuration differed con- siderably from that of the other three models. As mentioned previously, the values of Cq-ase and ( ! D i n t for configuration 4 are assumed to apply to the other ducted models.
for each of the four configurations. Figure 1.7 presents ‘Dext Configuration 2 with the smaller canopy and sharper nose and inlet lip had values of CD which were 0.002 lower than those of configuration 1 ext between Mach numbers of 1.05 and 1.28. These modifications resulted in a decrease in C D , , ~ from 0.044 to 0.042 at M = 1.05 and from 0.042 to 0.040 at M = 1.28. There was no change in the drag-rise Mach number which was 0.93 (based on dCD/dM = 0.10) in both cases. The subsonic- drag level in both cases was 0.017.
When the fuselage of configuration 2 was recontoured (based on an area-rule application at M = 1.2), a large drag reduction was achieved throughout the supersonic range of the tests. Both configurations 3 and 4 show CD = 0.035 between Mach numbers of 1.05 and approxi- ext mately 1 . 7 . Subsonic-drag levels of configurations 3 and 4 were 0.015 The drag-rise Mach numbers of both of these and 0.017, respectively.
= * configurations were again 0.93. 3 -
. - -
The data presented in figure 17 show that the redesign of the nose and canopy reduces C D ~ ~ by about 5 percent at M = 1 . 2 8 . By recon- tourlng the fuselage with only small changes in the maximum cross- sectional area, an additional 12-percent decrease is realized at the same Mach number.
CONCLUSIONS Results of minimum-drag tests of four specific versions of a swept- wing fighter-type airplane indicate the following conclusions : 1 . The configuration with the modified forebody (smaller canopy, sharper nose, and inlet lip) showed reduced values of external-drag coefficient at l o w supersonic Mach numbers. The modifications decreased the external-drag coefficient f r o m 0.044 to 0 . 0 4 2 at a Mach number of 1.05 and from 0.042 to 0 . 0 4 0 at a Mach nuniber of 1 . 2 8 .
2 . When the fuselage of the modified configuration was recontoured for an area-rule application at Mach number 1 . 2 , the external-drag 0.035 between Mach numbers of 1.05 coefficient was further reduced to and 1 . 7 1 .
3 . The drag-rise Mach number for each configuration was 0.93.
Langley Aeronautical Laboratory, National Advisory Committee for Aeronautics, Langley Field, Va., May 1 1 , 1956.
i 1. Whitcomb, Richard T . : A Study of the Zero-Lift Drag-Rise Character- istics of Wing-Body Combinations Near the Speed of Sound. NACA RM ~ 5 2 ~ 0 8 , 1952.
2. Holdaway, George H . : Comparison Of Theoretical and Experimental Zero- Lift Drag-Rise Characteristics Of Wing-Body-Tail Combinations Near the Speed Of Sound. NACA RM A53Hl7, 1953.
3 . Wallskog, Harvey A . , and Hart, Roger G.: Investigation of the Drag of Blunt-Nosed Bodies of Revolution in Free Flight at Mach Numbers From 0.6 to 2.3. NACA RM L53D14a, 1953.
4. Mitcham, Grady L., and Blanchard, Willard S . , Jr.: Low-Lift Drag and Stability Data From Rocket Models of a Modified-Delta-Wing Airplane With and Without External Stores at Mach Numbers From 0.8 to 1.36.
NACA RM L53A27, 1933.
5. Sears, R. I . , and Merlet, C. F . : Flight Determination of the Drag and Pressure Recovery of an NACA 1-40-250 Nose Inlet at Mach Numbers From 0.9 to 1 . 8 . NACA TN 3218, 1955.(Supersedes NACA RM ~50~18.)
I TABLE I I DIMENSIONAL AND MASS CHARACTERISTICS OF CONFIGURATIONS .
Wing:
Total area (excluding chord extensions), sq ft . . . . . . . . 4.53
Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . 3.40
Mean aerodynamic chord (excluding chord extensions) . . . . . 1 . 2 9
Incidence angle, deg . . . . . . . . . . . . . . . . . . . . . -I
Dihedralangle, deg . . . . . . . . . . . . . . . . . . . . . -5
Sweepback (quarter-chord line), deg . . . . . . . . . . . . . 42
Airfoil section at root, parallel to
free-stream direction . . . . . . . . . . . . . . . . NACA 65~006
Airfoil setion at tip, parallel to
free-stream direction . . . . . . . . . . . . . . . . NACA 65~005
Taper ratio . . . . . . . . . . . . . . . . . . . . . .
. . . 0.25
span, ft.. . . . . . . . . . . . . . . . . . . . . . . . . . 3 . 9 2
Vertical tail (extended to model center line and : not including dorsal fin)
Area, s q f t - . . . . . . . . . . . . . . . . . . . . . . . . 1 . 1 9
Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . 1 . 5 0
. . . . . . . . . . . .
Sweepback ( quarter-chord line), deg
Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . 0 . 2 6
Span, ft . . . . . . . . . . . . . . . . . . . . . . . . . . 1.33
Airfoil section at tip . . . . . . . . . . . . . . . . NACA 65~od.c
Airfoil section, 3.02 inches above
fiselage center line . . . . . . . . . . . . . . . . NACA 6 5 ~ 0 0 6
Configurations 1 Configurations 3
and 2 I and 4
I
Horizontal tail:
Total area, sq f't . . . . . . . . 1.28 1 . 1 4
Aspect ratio . . . . . . . . . . .
3 -5 3 -5
Incidence angle, deg . . . . . . . 0 0
Dihedral angle, deg . . . . . . .
5 -4 5 0 4 Sweepback (quarter-chord line), deg . . . . . . . . . . .
45 45 Airfoil section at root, parallel
to free-stream direction . . . . NACA 65~006 NACA 6 5 ~ 0 0 6
Airfoil section at tip, parallel NACA 6 5 ~ 0 0 4 NACA 6 5 ~ 0 0 4 to free-stream direction . . . .
Taper ratio . . . . . . . . . . . 0.15 0.15
2.12
span, f t . . . . . . . . . . . . . 1 =99
~~ ........................
. . . . . . . . . . .
. . . . . . . . . . .
. . . . . . . . . . .
a , . . . . . .
P - * M * -P 9-1 d A h . . .
. .
P 09-1
e * $
m e r( G4 r( k m a k m a k a W a 4J
s
r4
x
L
i : . 1
z
0 . 1 .2 . 4 . 7 .0 * 9 1 . 0 1 . 1 (a) Equivalent body (complete model).
.016 .012
5 .008
0 . 1 .2 . 3 . 4 - 5 .6 . 7 . 8 .9 1 . 0 1 . . 1 X / l (b ) Normal-cros s- sectional-area distribution.
Figure 2.- Equivalent body and normal cross-sectional area Of configuration 1 .
Model . 1 .9 1 . 0 1 . 1 0 *1 .2 . 3 . 4 - 5 .6 -7 . 8 x/z (a) Equivalent body (complete model).
(b) Normal-cross-sectional-area distribution.
.
Figure 3 . - Equivalent body and normal cross-sectional area of conf'iguration 2.
e. e.. e .e* e e. e. e e e e.. e.
e . . e . . e . e .
e . . e *
I 16 : : *: : UACA RM L56E25a
e . e .
- " e. e.. e e e 0 . e. e e .e. e. e.. e.
Model (a) Equivalent body (complete model).
.016 .012 nI N
. ook
n " .6 .9 1.0 1 . 1 0 . 1 .2 .5 .4 . 5 X I 2 ( b ) Normal-cross-sectional-area distribution.
Figure 4.- Equivalent body and normal cross-secticnal area of configuration 3 .
.
Model . 1 0 N
z
O O . 1 .2 .3 .6 .7 . 8 . 9 1 .o 1.1 X A (a) Equivalent body (complete model).
.016 .012 ru ry .ooe
. 2
selage (including d u c t ) .OOh Vertical t a 0 .1 .2 .3 . 4 - 5 .6 .7 .8 . 9 1 . 0 1 . 1 .It (b ) Normal- cross - sectional -are a d i s t r i b u t i on.
Figure 5.- Equivalent body and normal cross-sectional area of configuration 4.
NACA RM L56E25a 1 8 I .
C rl
-
N t - - - o o o o o m o m o m c 'i I C e o L e .me e o .........
. . . . : - **: *:i NACA RM L56E25a
........
- . . . . . . . . . . . . . . . . .
.........................
-- .. 0 . e . . e.. 0 . 0 . .
........................
........................
I
L-89 540 . 1
Figure 11.- Model-booster combination p r i o r t o launching.
' . .
. -
.
d -4J k a 3 k ry .
I I u'r J7 N3 a 3 n
%
I al m L d P I !
v3 rl d 0 d 9 9 I a ..
- - 27 0 . s o . 0 0 - 0 0 0-0 0.- 0. 0 0 i 0.0 0.
.
I I * .
r- e ri I n -P d E: ‘u, I d E- d d .
l-i
a
NACA - Langley Field, Va.