Graham, D. J.

English

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CELWTERISTICS OF A ! I ! R f A m m m E D  AIRCRAFT 
. :  I - PIBFORMANCE DATA 
."  By Donald J. Graham 
. , Amss Aejronautical Laboratory 
- - 
Approximately a pear ego a research program was formulated a t  
the Ames.Aeyonautica1 Laboratory aimsd a t  investigating thq possi- 
b i l i t i e s  of employing a wing of low aspect r a t i o  ar?b tr iangular plan 
form on a Qansonia ol- mdel*ately supersonic a i r c r a f t .  A wing w a s  
selected t o  be inveetigated concurrently i n  the subsonic, transonic, 
and supersoni'c wind-tamel f a c i l i t i e s  of the laboratory and, i n  
adaition, a t ' t r a l son ic  speeds by mans of the NACA wing-flow method. 
It was planned t o  determine thereby the effects  of wide variations i n  
both Reynolds n~mber and Yich number upon the character is t ics  of 
the subject configuration. 
I 
The ckoice of wing waa made on the basis of the bes t  exis t ing 
predictions of the pressure-drag churacter is t ics  of triawplm 
a i r f o i l s  i n  the moderately supersonic-speed region. The wing was 
of .5-perynt.-chordWthick symmtrical double-wqdge section with 
maximum thickness at 20 percent of the a i r f o i l  chord and had an 
aspect r a t i o  of 2 with a vertex anglo of 53'. The sweep of the 
leading edge thus amounted to  approximately 63O. 
I n  figure 1 i s  pictured the model, which was tested i n  the 
Amsa G by 3$ -foot tunnel tud the 1- by +foot supersonic tunnel, 
and the smallest scale m d e l  tes ted ,  - .w ing  was mounted on a 
slender cyl indrical  body which w a s  sting supported from the r ea r .  
. . 
Figure 2 is  a photograph of the model i n  the Ames 12-foot 
low-turbulence pressure tunnel and ehms the semispan configuration 
munted on a turntable i n  the tunnel f loo r .  
Aerodynamic characteristics of t h i s  wing were determined fo r  
Mach numbers from 0.1 to  1.5 and for  Reynolds numbers from 0.7 X 10 6 
to 27 x 106. ' W ~eynoMa nhib$variation &s confined t o  the . 
subsonic t e s t s ,  a l l  of the supersonic tests having been made fo r  
6 .  
-Reynolds numbere of the order of 1 x 10 . 
A considerable port  ion of the r e su l t s  of t h i s  i n v e ~ t i g a t i o n  
will be published shortly.  The object of the present paper is to 
summarize the principal results which are involved i n  a prediction 
of the performance a d  the s t a b i l i t y  and control c ~ q c t e r i s t l c s  
of a low-aspec'tratid t r i a n g u l h n g  a i r c r a f t .  , ' 
https://ntrs.nasa.gov/search.jsp?R=19790079939 2020-03-21T20:13:43+00:00Z
& variation of the miaaum drag ~oefficlent with &oh number 
. f o r  the t r imgu la r  wing is  shown i n  fi6txt-6 3 .  It will be noted - 
t h a t  the r e su l t s  from the hues 1- by .3L-foot tunpel (unpublished 
2 - .  
data) presented f o r  Mach nxnbsrs from 6.5 to  1.5, fndicated by the 
so l id  Urn, appear t o  remonabky b.ridgq the ~ a p  between the lo? 
.< . t  
speed value from the Ams 7- by 10-foot twklel (reference 1) arid the - b t  - 1  , 
value f o r  a Mach ~ rmber  of 1.5 f r o 3  the Amas 1- by +foot supersonic 
t-el (refei'ence 2 ) .  The kigher Reynolds number data from-the 
Amee 12-foot low-turbulence preseure tunnel (reference 3)  are not 
i n  such closa agreement with the high subeonic Mach n i i e r  data  
*om the Ams 1- b j  %-foot tunnel aa could be desired despf t e  
' allowance f o r  the difference i n  acale. '1t g3hoGd be emphaoized, 
however, that the respective t e s t  conditfone were dissimilar. The 
wing i n  the Ames -1- by 31-foot t m e l  wae mounted on a thin body, 
- 2- , 
the drag of which could not  readily be s e p p t e d  from that' of the 
combination; whereas, the data  from the Ames 12-foot low-turbulence 
pressure t m a l  shown are for  the w i q  alane. The ef fec t  of adding 
a fuselage to  the mod31 wing i n  the Amss 12-foot low~urbulence  . 
pressure tunnel was to' displace the curve of miliimm dreg coefficient 
above t h a t  of the wing i n  the Am8 1- by +-foot tunnel.. no 88tiS- 
2 
factory explanation h ~ s  ye t  been forthcoming f o r  the seemingly ear ly  
r i s e  i n  the drag coefficient with Mach nuribsr evidenced by the r e su l t s  
from the h e 3  12-foot law-turbulence pressure tunnel. 
ALeo %how f o r  coxup&ison i n  figure -3 are  minimum drag coef- 
f i cients fo r  a Gpercent-chord-thick symmetrical double-wedge 
aii-f'oil section f r o m  two-dimensional t e s t s  i n  the Ames 1- by 31-foot 
2 
tunnel. The favorable e f f ec t s  of sweep and aspect-ratio reduction 
axe apparsm here. 
* 
It was Inferred a t  the beginning of the paper tha t  the prosent 
wing was selected because, from theoret ical  considerations, it had 
the lowest pressure drag f o r  the practicable thickness d i sk ibu t ions  
of the given triangular plan form at  moderately supersonic speede. 
Subsequent tests', however, In the Am6 1- by +foot supersonic 
tunnel and the Ames 1-.by 3*--foot tupnel shoved lover actual minimum 
C 
d r q  coefficient8 a t  a Mach n&er of 1.5 f o r  a wing of the same plan 
form with the maximum thickness a t  50 percent of the a i r f o i l  chord, 
an ef fec t - t raced  t o  the differences i n  tip f r ic t i ,oa  drag of the two 
surfaces. Eence, i f  q y  usefhl flrnction such as e t ruc tura l  con- - . 
venience we:w to be served by locating tpe'maximum thicisness i n  the 
v ic in i ty  0.f the midchord, -re would apparently be no assoc iabd  
penalty i n  minimum drag. 
In f igure 4 the variation of maxl.mm l i f t d r a g  r a t i o  wfth Mach 
number i s  presented. The diffeiwnces i n  the subsonic-speed charac-- 
t e r i s t j c s  as determined fn the va-ioue f a c i U t l e s  appear to be 
consistent with the co&egonding differences i n  the Reynolds 
auxibers of the respect i re  t e s t s ,  .me eubsonic-speed l i f t cd rag  
ra t ios ,  althou& seemingly lwr,  d g h t  mesonably be expected t o  
improve s o m v h t  wfth increasing Reyuolde nurcbers, a s  was observed 
i n  the case of the subsonic speed character is t ics ,  
F u r t h e m r e ,  f o r  theso t e s t s ,  the w i ~ g  had sharp leading edges 
and, hence, did not rea l ize  an appreciable m u n t  of the possible 
leadinq-edge suctioq which would fur ther  boost the maxim l i f t -  
drag r a t i o s  i n  the speed range under consideration. 
., 
Previously reported t e s t s  (reference 4) i n  the Ames 1- by 3-foot 
supersonic tunnol a t  a Mach number of 1.3 TJith the leading edges of 
t h i s  wing rounded have indicated the attai~llhent of a s i m i f i c a n t  
but by no mans najor portion of the theoret ical  leading+dge 
suction. Figure 5 ,  the material f o r  ;rhich wads presentgd a t  the 
NACA Conference on Su?orsonic Aerodynanics a t  the Langley - 
Laboratory, June l9--2O, 1947, i l h s t r a t e s  the var iat ion of l i f t -  
drag r a t i o  with l i f t  coofficient at'a Mach number of  1.5 fo r  the 
w i n g  with sharp leading edge and with the loading edtp rounded to  
approximate the noee radius of a 3-percent-chord-thick NACA 6 w e r i e e  
a i r f o i l .  Rounding the leading edge, while ra i s ing  the ;naximm l i f t  
drag r a t i o  by decreasing the drag due to  l i f t ,  did not afYect the 
minimum drag. These r e su l t s  should not be taken as evidence of the 
maximum gain t o  be expected from leading-edge shape mdi f i ca t ion  
because the'eubject wing section was not selected with t h f s  objective 
i n  mind. It appears l i k e l y  t h t  a t  fuU=-scale Reynolds numbers the 
use of a i r f o i l  sections with,rounded nose contours of the subsonic 
type on wings w i t h  highly swept leading edges would, by real iz ing a 
greater par t  of the possible leading-edge suction, afford con-.  
eiderably higher maximum lift-drag r a t i o s  ' a t  low supersonic Mach 
numbers than those indicated i n  figures 4 and 5 .  . 
An additional f a c t  of in t e re s t  is t h a t  the l i f t  coefficients 
corresponding to the maxirnunliftrdrag r a t i o s  were found t o  be 
sensibly independent of Mach number, having varied but inappreciably 
f r o m  a'value of about 0.2 over the range of the tests. 
, 
The slope of the l i f t  curve of the triangular wing as  a function 
of Mach number i s  shown i n  figure 6. Sat isfactory agreement i s  
evident both between the r e su l t s  of the various wind-tunnel t e s t s  
and the calculated subeonic and supersonic values. The variat ion 
with Mach number is regular and apparently freo from abrupt dis- 
continuities a t  t r a n ~ o n i c  speeds . 
Froh the standpoint of perfqmce at transonic Mach numbers 
the r e su l t s  of reaoarch to  date i n d i c a t ~  the low-aspectrratio 
triangular wing to  be a practicable l i f t i q g  mrface f o r  a ahor* 
range luterceptor a i r c ra f t .  Were a v t n g  section t o  be selected a t  
t h i s  dase fo r  an a i r c r a f t  designed t o  f l y  at transonic o r  mderately 
euprsonic  Mash nuribera w i t h  tke type of w i n g  plan form undpr .d iy  . . . .  L .  . .  - 
cusoion, a profile having the general shape of t3e RACA 64-eriee 
o r  G-eries a i r f o i l  ssctions would bs recommeaded because of the 
higher m a x i m u  l i f t -drag  r a t ioe  afforded. / 
1. Roee, Leonard M. : Low-Speed Dvost igat ion of a Smll Trientp3a.r 
Wing of Aspect Ratio 2. I - Effect  of Combination with a 
Body of Revolution and Height above a Ground. Plane. NACA 
RM No. AZ03 ,  lgIc7. 
- 
2. Vincenti , Walter G . , ~ i e l s e n ,  Jack N . ; and Matteson, Frederick H . : 
Inveetigation of W i n g  Characteristic8 a t  a Mach Number of 1.53. 
I - Tr1ar;gulas Wings of Aspect Ratio 2. NACA RM No. A7110, 
1947 
, , -  
3. ~ & a r d s ,  George 6 . ,  and Stephenson, Jack D . : Tests of a Tri-  
angular W i n g  of Aspect Ratio 2 i n  the Ames 12-Foot Preseure 
Wind T~nslel. I - Effect of Reynolds Number and Mach Number 
>on the Aerodynamic Characteristics of the Wing with 
Undeflected Flap. NACA Rm No. AVO?, 1947. 
4. ~ h c e ~ t i ,  Waltor G. : Emerimwtal  Characteristics of F i n i t s  
Span Winge. NACA Conference on Supersonic Aerodynamice, 
June 19-20, 1947. 
Figure 1.- Model tested in the Arnes 1- by 3 ;-foot tunnel. 
Figure 2.- Model tested in the Ames 12-foot low-turbulence 
pressure tunnel. 
MACH NUMBER, M 
Figure 3. - The variation with Mach number of minimum drag coefficient 
for triangular wing. 
L / ~  
MAX. 
MACH NUMBER, M 
Figure 4. - The variation of Mach number with maximum Lift-drae ratio 
LE. RAD. = I X CHORD \ 
Figure 5.- The effect of leading-edge radius upon Wt-drag ratio of 
triangular wing at a Mach number of 1.53 and a Reynolds number 
of 1 x lo6. 
MACH NUMBER, M 
Figure 6. - The variation of Mach number with lift-curve slope of the 
triangular wing. 
r 7 4 ~ )  0 


Characteristics of a triangular-winged aircraft. 1: Performance data

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