Correlation of the Drag Characteristics of a P-51B Airplane Obtained from High-Speed Wind ...

32 Pages·2005·9.66 MB·English

ACRNo. 4K02
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
February 1945 as
Advance Confidential Report 4K02
CORRELATION OF THE DRAG CHARACTERISTICS OF A P-51B AIRPLANE
OBTAINED FROM HIGH-SPEED WIND-TUNNEL AND FLIGHT TESTS '
By James M. Nissen, Burnett L. Gadeberg, and
William T. Hamilton
Ames Aeronautical Laboratory
Moffett Field, California
,
WASHINGTON
it ' NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of
advance research results to an authorized group requiring them for the war effort. They were pre-
viously held under a security status but are now unclassified. Some of these reports were not tech-
nically edited. All have been reproduced without change in order to expedite general distribution.
A-62
, - MACA RCB Bo. 4K02
NATIONAL ADVISORY COMMITTEE FOR AEBOBAUTICS
ADVANCE COYFIDEYTSAL REPOIZT
-._-
COBRELATIOY OF THE D U G CSARACTXRISTICS QE A P-51B AIRPLAYE
OBTAINED F2.02~1H IGZ-SPEED VIXI-T3NNEL AYD FLIGHT T3ST.S
By James M. Nissen, Burnett Lo Gadeberg, and
W i l l i a m T. Xamilton
"Editorial Wote
"Wiih f u l l recognition of the hazards in-
volved, the senior antho? of this report, N a C l
. Test Pilot Janes K. ;;issen macle a series o f
dives with the Y-513 airplane without propeller,
because the need.ed data could be obtained i n no
other way. After three siiccessful dives at s'ic-
cessivel;; higher Psch nun3ers, the towrope con-
nection broke at a l o w altitude and the towrose
wrapped around his nirplsne. J i t h great sr;ill
and courage Mr. Wissen stayed with the airplqne
and made a forced landing i n rough terrsirL witl1-
out damaging the elaborate researcn instrumenta-
tion. Although the airplane W ~ Swa shed O U ~ ,: Ir.
Nissen escaped with minor injuries,
In order t o obtain a correlation of drar dstn from wind-
tunnel and f l i g h t t e s t s at high Mach nurbers, a P-513 air-
plane, with the propeller removed, was tested i n f l i g h t at
Mach numbers up t o 0,755, and the results were c o ~ p a r o dw ith
wind-tunnel t e s t s of a 1/3-scal~m od-1 of the sirplane.
.' The t e s t results show that the c?raP; characteristics O f
the P-51B airplane can be predicted with sztisfactor;T acc'i-
racy from tests i n the 16-foot uind tunnel of the Ames l e r o -
I nautical Laboratory at both high an3 low Mach ni-mbers. I t i s
I -' considered that t h i s resuit is riot unique with .chis airplane.
NACA ACR No. 4K02 2
INTRODUCTI O??
Practica.lly a l l the available data on the drag chsrac-
t e r i s t i c s of ai.r.r,lanss at hfgh spee6.s have been obtained
from wind-tunnel tescs, The reliability of these data has
been questioned becauze of the fact that (l! sui*veys have
indicated that the us-lial strut-support systems nave a marked
infl-uence on the flow at the position of the model, (2) the
calculated wall interference, alvays R soinewhat dubious fac-
t o r , increases ranidly with Mach nunber, and (3) the magni-
tude of the effects o f Reynolds nuuiber at high speeds i s un-
known, I t was evident that a c0111pa.rison of accurate f l i g h t
and wind-tunnel t e s t data “15s needed t o deterI;ine t h e relia-
b i l i t y of the wind-tunnel t e s t data.
Of the limited amount o f flight data avail.tble fro&
high-speed dives, none were considered. satisfactory for such
a comparison, For the most part, the measurcxents of air-
speed and altitude vere not above suspicion. nh ven f o r those
cases wherein t h i s objecti3n co-ild not be raised, the prob-
able error introduced i n attem?’iing t o correct fdr the in-
fluence of the propulsion systein i n tlie deterrfiination of
drag made comparisons with high-speed wina-tunnel data f o r
propellerless models of doubtful valii.e,
The purpose of the investiqation herein dlsc’isse; was
t o obtain an accurate co~noarisono f the drag conffisients at
high Mach numbers as measured i n fii,:ht and i n the vind tiin-
nei, In order t o obtain flight nata s t r i c t l y c3nlp~rablet o
those obtained from a model of the P-Ell? airplzne i n the
h e s 16-foot sind txnnel, 5he f l i g h t a a t s vFre o5tnined i n
dives of tnis airplane wit:; propeller rern3vc.d.
It
Description of the hirplane
The Xorth huierisan P-51B a’rp~ane use?. for the f l i g h t
t e s t s was a single-en<ine, low-wing, cantilever nononlane
with retractable landing gear and parLial-span plain flaps
Figure 1 i s a three-view drawinp of the airpla2ep and tk?
p h o ~ o ~ r a pohf figure 2 shows the airplane as instruaentecl
f o r the f l i g h t t e s t s ,
NACA ACR No. 4K02 3
The pertinent specifications of the airplane as tested
were as follows:
Airplane . . . . . . . . . .
Manufacturer Nortli American Av4 ation,
. . . . . . . . . . . . . . . . . . . . Inc.
Type P-SiE-L-'T-4
. . . . . . . . . . . . . . . . .
Army number 43-i%11
Weight
. . . . . . . . . . . . . . . . . . . .
Ell19ty . . . . . . . . . . . . . . . . . 7503 lb
Normal gross . . . . . . . . . . . 97G3 l b
As flown f o r these t e s t s 7335 Ib
Wing loading
. . . . . . . . . . . . .
Normal gross . . . 35.9 lb per sq f t
As flown for these t e s t s 31,4 lb per sq f t
a I
Center of gravity (as flown for these
. . . . . . . . . . . . . . .
t e s t s ) 25.3 pgrcent N.A.C.
Wing
. . . . . . . . . . . . . . . . . . . . .
S p a n . 37.03 f t
I - . . . . . . . . . . . . . . . . . . . .
Area . . . . . . . . . . . . . . . . . .23.3 ,2 sn_ f t
Root chord 164 ic.
. . . . . . . . . . .
I ". Airfoil sectio.n. . . . . . . . . . . . .Nd.A-.XA.Ch. l.ow drag
Aspect ratio 5.815
. . . . . . . . . . . . . . . . . . .
Taper ra.tio 0,449
Thickness r a t i o (measiired)
. . . . . . . . . . . . . . . . . . . . .
Root. . . . . . . . . . . . . . . . . . . . . 0,163
T i p . . 0.107
Incidence
. . . . . . . . . . . . . . . . . . . . .
Root 1OOG'
. . . . . . . . . . . . . . . . . . . .
T i p . . -9'53'
. . . . . . . . .
D i h e d r a1 ( 2 5 - p e r c e n t .. ch ord 1i n e ) 5 O
SweeFback (measured alon.g. le.ad.in.g .ed.ge.) . ,. . . . 3 O35'32"
Mean aerodynanic chorc? 79,G in.
Location o.f .m.ean. a.er.od.yn.am.ic. . . .
chord Leading edge of I!.A.C.
i s 6.1 in. aft and ci.2
in. above the leading
edge of the wing r o o t ,
Horizontal tail surfaces
Stabilizer
. . . . . . . . . . . . .
Area 27, 5t? sq- f t (incl-.idjnq
. . . . . . . . . . . . . . . . . fuseleqe)
Span . . . . . 13 f t 2.12 in.
Incidence (for these tests) 1' (relative t o
fuselage axis)
Elevators ( a l l metal f o r these t e s t s )
. . . . . . . . . . . . . .
Area 13,05 sa_ f t ( t o t a l )
. . . . . . . . . . . . . . . . .
Span 13 f t 2.12 in.
NACA ACR No. 4K02 4
Vertical tail surfaces
Fin . . . . . . . . . . . . . . . . . . .
Area . . . . . . . . . . . . . . . . 9.61 Sq f t
S p a n . . 5'7 in.
. . . . . . . . . . . . ,, . . . . .
Setting lo l e f t
Rudder
. . . . . . . . . . . . . . . . . . .
Area . . . . . . . . . . . . . . . . . 1 0.41 sq f t
S p a n . . 74,4 in.
The stabilizer incidence was set at 1' instead of the
normal 2' for the P-51B-1-NA airplane, i n order t o reduce
the elevator angle required for t r i n at high Mach numbers,
and a pair of metal-covered elevators was substituted for
the usual fabric-covered elevators ,.
I n ordar. t o simulate as closely as possible the model
as tested i n the wind tunnel, the propeller was removed. A
spinner w a s installed t o preserve the smooth air flow over
the forward portion of the fuselage. The tow-release mecha-
nism was housed within the spinner and was f i t t e d flush with
the soinner nose (fig. 3). The release mechanism was mechan-
i c a l l y operated by the pilot. A special hydraulic pump,
e l e c t r i c motor, and batteries were installed t o activate the
landing flaps and gear, The carburetor a i r scoop was sealed
about 3 feet from the scoop l i p , the bomb racks were removed,
and the surface of the airplane sanded w i t n fine sandpaper,
shellacked, and waxed,
Instrument ati on
Standard NACA photographically recording instruments
were used t o obtain airspeed, altitude, and normal and longi-
tudinal acceleration as a function of time. Two complete
and independent sets of instruments were installed for the
f l i g h t t e s t s . Each system of instrumentation u t i l i z e d , as
sources of s t a t i c and t o t a l pressures, a freely swiveling
Pitot-static heai?. These two p i t o t - s t a t i c heads were
mounted on booms located beneath and extending approximately
0.8 Of the local wing chord ahead of each wing tip. (Cf,
figs. 1 and 2.) A service total-head tube of round section,
which was used i n conjunction with fuselage static-pressure
orifices, was mounted beneath the right wing on the standard
airspeed mast.
The pressure lines from the Pitot-static heads t o the
recording instruments were made as s3ort as possible t o
NBCA ACR No. 4K02 5
minimize lag, and the lines t o the record.ing airspeed meters
were balanced s o as io give equal flcw rates i n the static-
and total-pressure tubes. Each pitot-static head consisbed
of two static-pressure tubes and one total.-head tube, which
permitted the u s e of independent sources of s t a t i c pressure
f o r both the airspeed and altitude racorders. Ground t e s t s
of a rrock-up of the airspeed and alti.tude pressure lines in-
dicated that the lag i n the system, at the rfiaximum rates of
descent, caused an error i n the recorded altitude of only
250 feet.
The recording instruments, as installed i n the airplane,
could be read t o *2 miles per hour for the airspeed, *250
feet f o r the altitude, i O o , O l g for the longj tudinal accelera-
tion, and *O.lg f c r the aormal accelaraticn,
The recording acceierometer, from tdhich the drag data
were deternined? was mounted 4 feet aft and 1 foot above the
center of gravicy of the airplane, ?'he crfcc'S of angular
accelerations of the air2lane during ike Ci~vt?o~n the re-
cordings of the ~ c ~ ~ i e r o i n ewtaes r found t o be negligible.
Calibration of the Pitot-static Tubes
A correction €or the position error of the Pitot-static
tubes was, detzrmined by flying the airplane at a known con-
stant pressure altitude at various airspeeds, while records
were made of the airspeed and altitude, It was assuned that
the measurements o f the t o t a l pressure wcre cc~rrecta nd that
the variation of recorded altitude with airspeed at the con-
stant pressure altitude resulted from the position of the
s t a t i c tubes. Assuming the error i n the measured s t a t i c
pressure t o be the complete error, the indicated airspeed
was corrected by the use of the standard formula
where
Vi correct indicated airspeed, miles per hour
8 free-stream t o t a l pressure, inches of mercury
p free-stream static pressure, inches of mercury
NACA -4CR No. 4K02 6
This correction i s presented in figure 4 and was applied
t o the airspeed calibration. Since the rcaximum error of al-
titude, as determined by t h i s calibration, was smaller than
the l e a s t reading of the altimeter, no attempt was made t o
correct the altimeter readings f o r position error. The ac-
curacy of the swiveling pitot-static head has been investi-
gated at Mach numbers up t o 0.80 in the 16-foot wind tunnel,
and the results showed the effects of compressibility t o be
negligible over the flight range investigated.
Tests
In order t o determine the drag coefficient of the air-
plane at high Mach numbers i n a configuration that would
lend i t s e l f t o direct correlation with wind-tunnel t e s t s ,
the airplane (without propeller) was towed t o high altitudes
by a Northrop P-61A airplane (fig. 5), where the pilot Of
the P-51B airplane released the tow. The airDlane was then
dived t o high Mach numbers and a t the completion of the dive
i t was landed on the surface of a dry lake.
In order t o obtain the high Mach numbers at a safe a l t i -
tude, the airplane was towed t o as high an altitude as pos-
sible, which was approximately 28,000 f e e t pressure altitude
for the third f l i g h t i n which a Mach number of 0.?55 was ob-
tained.
During the dives the radiator-scoop flap was locked i n
the flush position at all times.
Three dives were made successfully, each t o successively
higher Mach numbers, but on the fourth attempt a forced land-
ing was necessitated soon after take-off due t o an unex-
plained, prematurs release of the tow cable from the tow
plane. The forced landing damaged the P-51B airplane beyond
repair, and hence terminated t h i s set of tests.
Computation of the Drag Coefficient
The drag coefficient was computed from values of the
airspeed, altitude, longitudinal acceleration, and normal
acceleration by the use of the following equation:
-
CD = [AZ s i n a AX cos a]
9
NACA ACB Eo. 4x02 7
where
CD air7lane drag coefficient
W airplane weight, pounds
S wing area, square feet
q dynamic pressure, 2ou.nds per square foot
AZ algebraic sun of ooaponents, along airplane Z-axis, o f
airplane accol erntion and. acceleration due t o gravity,
i n terms of stanilard gravitational unit (32,,2 ft/secz).
Positive when dircctqa up'.rard as i n nornsl level
f 1i ght
AX algebraic sun of components, along airplane X-axis. of
airplane acceleration and accslcra.tion due t o gravity,
i n terms of standard gravitational unit (32.2 ft/scc").
Positive when directed forwa.rd as i n 8. take-off.
n angle of attack o f reference l i n e of the accelerometer,
degrees
The angle of attack of the airplane f o r a given l i f t
coefficient was determined froni measurerents made of a P-51B
airplane i n the Langley full-scale wicd tunnel. !TO effects
of compressibility on this determination were considered.
An attempt was made t o determine the drag coefficient
Of the airplane by the use of time histories of altitude and
airspeed and by use of the following equation:
1
dh /d t dv /d t
V
E?
where
dh/dt r a t e of change of true altitude, feet per second
dv/dt r a t e of change of trne airspeed, feet per second
per second
g acceleration of gravity, 32.2 feet per second per
second
It was found; however, that the slightest inacmracy in
the determination of the airspeed curve and i t s slope would
NACA ACB No. 4K02 8
produce absurd results. Negative drzg ccefficients could be
calculated by t h i s method. as well as excsssively high drag
coefficients. Consequently, this method was discarded and
a l l flight-drag results presented in t h i s report have been
derived from accelerometer data.
Plight- Tes t Data
Dives were made t o 1:ach numbers of 0.710, 0.730, and
0.755. Before each f l i g h t , the surface of tho airplano was
carefully wiped clean t o preserve as smooth a finish as pos-
sible during the dives. However, due t~ the short length of
the oiiad s t r i p on the lake bed, which was used for take.-off
(about 2503 rt!, the airplane gntnprel:. some duct on the lead-
ing edges of :he wing and tail 3116 ?arts of the fuselage;
the airplanes ran off the end. of tlie oiled take-off
s t r i p arid onto the dusty surface of the lake bed-
Of the three flights made, the airplaxe :.tad the most
dust on i t s surface during the f i r c t flight (flight 1081,
the l e a s t during the s2ccnd flight (flight lOS)$ and an in-
termediate amount during the l a s t flight (flight 110). The
dustiness of the windshield, the nose section, and the lead-
ing edges of the wing and radiator scoop after the f i r s t
dive may be estimated from figures 6, 7, and 8. (The crosses
have been made by wiping the surface free of ci~lstw ith a
cloth,)
The results of the dive tests are presented i n figures
9 t o 1 2 , which show l i f t coefficient and drag coefficient
plotted as a function of Mach number f o r the three dives,
and i n figures 13, 14, and 15, which show the variation of
drag coefficient with l i f t coefficient o f the airplane at
Mach numbers below that of drag divergence (the Mach number
att which the drag characteristics diverge from their low-
speed trend as the Mach number is further incr?ase!L).
On figures 9: 10, and 11 faired curves have Seen drawn
through the t e s t points as well 2s a curve foilowing the
points. It i s apparent that cturing all three of the flights,
the drag coefficient varied with the l i f t coefficient at
Mach numbers both above and below that of drag divergence,
The faired curves on these figures were drawn a f t e r sonsider-
ing the variations of figures 13, 14, and 15,
A comparison of the data of figures 9, 10, and 11, as
shown i n figure 12, indicates that. the ninimun drag of the
NACA ACR No. 4K02 9
airplane was affected by the presence of dust on the surface
of the airplame and that, as would be expected, the dustier
the surface, the higher the minimum drag of the airplane.
It is noteworthy that the Mach number of d-rag divergence and
the variation of drag coefficient with Mach number above the
Mach number of drag divergence are essentially unaffected by
the presence of dust on the airplane,
It i s believed that the variation of drag coefficient
with l i f t coefficient at tne l o w values of l i f t coefficient,
shown on figures 9, 10, and 11, aay be due t o a fore-and-aft
movement of the transition l i n e of bo’undary-layer flow from
laminar t o turbulent flow. Such a movement of the transi-
tion point on the wing of the P-51B airplane i s possible, be-
cause i t s a i r f o i l section has a very small pressure gradient
at l i f t coefficients near i t s design value, and hence is
very c r i t i c a l t o surface waviness, which might well vary
with the load on the wing.
Few data were available from the dives t o show the var-
iation of drag coefficient with Reynolds number, but the
data that were available (below the Nach number of drag di-
vergence) seemed t o indicate very l i t t l e , i f anye variation
at constant l i f t coefficients. This is not at a l l conclu-
sive, but i t i s interesting i n the light of the resalts re-
ported i n reference 1.
WIND-TUNNEL INVEST IGAT I 3N
Description of Apparatus
The model t e s t s were cond~ctedi n the Ames 16-foot high-
speed wind tunnel. This wind tunnel i s of the single-return,
closed-throat type and has a circi~larc ross section through-
out i t s length, Two 5-percent-thick front struts and a sin-
gle 7-percent-thick rear strut supported the model during
the tests. (See fig. 16.) A l l three struts were unshielded
and had the transitions of their respective boundary layers,
from laminar t o turbulent flow’, fixed at their lo-percent-
chord points. With the model !iiounted i n the wind. tunnel,
t e s t Mach numbers as high as 0,825 were reached. The turbu-
lence level i n the 16-foot wind tunnel i s very low, approach-
ing closely that of wind tunnels designed especially t o have
low turbulence.