tests model of aspect ratio 4 - unt digital library/67531/metadc58274/m2/1/high_res… ·...

. .- . COPY

TESTS O F A MODEL HORIZONTAL TAIL OF ASPECT RATIO 4.5

IN THE AMES 12-FOOT PRESSURE WIND TUNNEL.

I - QUARTER-CHORD LINE SWEPT RACK 35O

By Bruce E. Tinling and Jerald K. Dickson

Ames Aeronautical Laboratory -

- Moffett Field, Calif.

-. -.

..

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

WASHINGTON September 9, 1949

UNCLASSIFIED

By Bruce E. T i n l i n g and Jerald K. Dickson

Wind-tunnel tests &ve been conducted to evaluate the independent effects of Reynolds and Mach numbers on the aerodynamic characteristics of a horizontal tai.1 of aspect ratio 4.5 equipped with a p l a k sealed elevator with a tab. The line j o i n ing the qyarter-chord points of the airfoil sections was swest back 35' and the thickness distribution normal to this lFne was the NACA 64AOlO.

The Reynolds nunib- was varied from 2,000,000 to 1l,OOO,OOO at a Mach number of 0.21, and the Mach nmiber was varied-from 0.21to 0.94 at a Reynolds number of 2,000,000. Lift, drag, pitching maat, elevator hinge moment, tab hinge moment, streamKise distribution of static pressure at the midsemispan, and. pressure difference across the elevatcn" nose seal were measured.

An increase of Reynolds number f r o m 2,000,000 to ll,OOO,OOO had little effect other than to increase the anglMf-Etttack rage over which the variation of lift with angle of attack w a ~ linear.

Brupt decreases in liftrcurve slope occurred at'a Mach number of about 0.93 asd in elevator lift effectiveness at a Mach nuniber of about 0.87. The Mach numbers at which marked changes in the elevator hinge moment coefficients occurred were depcmdent upon the magnitude of angle of attack and of elemtor deflection. In general, however, the changes of elevator hing-t cbefficient were gradual as the Mach nuniber was increased to 0.83. The tab wa8 effective throughout the N c h number rage. Calculations indicated that incorporation of sufficient sealed internal balance to reduce the variation of elevator hinge moment with elevator deflection by 50 percent at a Mach number of 0.21would cause only a -percent reduction f o r elevator deflections greater than 4O at a Mach number of 0.93.

r

2

INTRODUCTICN

NACA FM ~ 9 3 1 3

A systematic investigation af control-surface characteristics ha0 been undertaken at the Ames Aeronautical Laboratory to determine experi- mentally the cantrol-effectiveness and hfnge-mment parameters for comparison with those.predicted by liftiwurface theory. References 1 through 4 present results of low-speed wind-tunnel tests of both swept and unswept horizontal tails of several aspect ratios, all having the same taper ratio and airfoil section.

The tests reported herein were conducted to evaluate the effects of compressibility and dynamic scale on the control-urface characteristics of a horizontal t a u with 35O of sweepback. The lo-speed aerodynamic characteristics of a geometrically similar horizontal tail k v e been reported in reference 2. SFnce this model a b 0 represents a wing with a fUl1”span flap or devon, drag and pitchwme& data are included in addition ’ to lift and hing-ent data.

drag coefficient (3) elevator hing+m”t coefficient

tab hinge-moment coefficient tab hinge moment ( 2q MAt

lift coefficient (F) pitching-mment point of the

m e a n aeroaynamic chord

pressure coefficient

critical pressure coefficient, corresponding to a Mach munber of 1.0 in a direction perpendicular to the quarter- chord line of the a i r f o i l section

pressure coefficient acros6 the elerntopnose seal (pressure below the seal minus pressure above the seal divided by the freestream dynamic pressure)

XACA RM ~9313 3

P

.

Reynolds number

first moment of (3

the elevator area behind the hinge line about the hinge IFne, feet cubed

-

first m o m e n t of the tab area behind the tab binge l ine about the tab hinge Une, fee t caed

horizontal-tail area, s q w e f e e t

airspeed, feet per second

speed of sound, feet per second

semispan, measured perp&c~i@- t o the plane of symmetry, feet

chord, m e a s u r e d parallel t o the plane

mean aerodynamic chord 6213 , of symmetry, feet

feet

chord of the elevator behind the hinge line measured perpendicular t o the hinge line, feet

local s ta t ic pressure, pounds per squsre foot

f'ree-stream s ta t ic pressure, pun& per square foot

fre+stream aynamic pressure, pounds per square foot

la teral distance normal t o plane of symmetry, feet

corrected angle of attack, degrees

angle of attack, uncorrected for tunne1"Wall interference and angl-f-ttack counter correction, degrees

elevator deflection (poeitive t o fncrease lift) measured in * a plane normal t o the elevator hinge line, dewees

tab deflection (positive t o increase l i f t ) measured in a plane normal t o the tab hinge line, degrees

density of air, slugs per c a i c foot

absolute viscosity, slugs per foot second

4 NACA RM AgGl.3

(measured through G O ) , per degree

=(?J cLF6t=0 (measured through 6e=O), per degree

-e =@J dt* (measured through 6e=O), per degree

The subscripts outsidethe parentheses represent the factore held constant during the measurement of the parameters.

MODEL

The semispan model tested in t h i s hvestigation represented a horizontal t a i l of aspect ra t io 4.5 and taper ratio 0.5. The a i r fo i l section was the NACA 64A010 (table I) in planes inclined 35O to the plane of symnetry-(fig. 1). The quartez-chord line of the a i r fo i l s e c t i w was swept back 35O. This line W E a t 27.8 percent of the chord measured parallel t o the plane of symmetry. The t i p shape was formed by rotating the section parallel t o the updisturbed stream about a line inboard of the t i p a distance equal t o the msxlrmnn tLp ordinate.

The model was equipped xfth a full-span, radius-nose, sealed elevator, the chord of which was 30 percent of the chord of the a i r fo i l sections. The rat io of elevator area behind the hinge l ine t o the to ta l model area was 0.271. The e1evis;tor was attached b t h e stabilizer by hinges a t 34, 80, and 96 percent of the semispan. These hinges a& a c l o s e f i t t k g block a t the plane of symmetry divided the sealed balance chamber i n t o '

1

RACA RM ~ 5 ~ 1 3 5

three separate sections. The seals were fitted clogely to the ends of each section to reduce leakage to a min-tlmnn. The elevator was eguippd with &z1 unsealed tab, the area of which was 6.5 percent of the elevator area- and which extended frola 23.7 to 44.8 percent of the semispas. m e elevator and the tab gaps axe Shawn fn figure 1.

The stabilizer was constructed of solid st-eel and the elevator of alumhum alloy. The model vas mounted vertically with the wind-thnel floor serving as a reflection plane as shown in figure 2. The rotating turntable upon w s c h the model. was mounted' is directly connected to the forc-easuring arnratus. The elevator and tab hinge moments were measured with resietanc~tgpe electric strain gages. The elevator gage was beneath the turntable cover plates, rmd the tab gage wa8 contained within the elevator. The elevator deflection was remotely controlled and the tab deflectim was set by m e a g s of as indexing sggtem built into the tab asd elevator. 'The gap between the elevator Etnd the reflection plane was appro"te3.y 0.02 inch when the elevator was undeflected.

A streamxise raw of orifices vas provided at 50 percent 09 the semispan to measure the chordwiee distribution of static pressure. S i x orifices were located in the balance chamber, one on either side of the seal at 16, 48, and 90 percent of the semispan, to measure the pressure differences across the elevsto~ose seal.

The data have been corrected far the effects of tunne1-WaJ.l inte?+ ference, far constriction due to the presence of the tunnel walls, and for model-support tare forces.

Tunnel-Hall Interference

Correctians.to the data for the effects of t u n n e l - interference have been evaluated by the methods of reference 5 , using the theoretical span loading calculatedby the methods of'reference 6. The corrections added to the drag and to the -le of attack were:

-hcL = 0.329 CL, degrees

At3D = 0.00502 CL' No attempt was made to separate the tuqel-wall interference effects

1 resulting f r o m lift due to elevator deflection and lift due to angle of attack. No corrections were applied to the p i t c h i n m a t or hing- m m e n t data. . .

c

* 6

Constriction Effects

NACA RM ~ 9 3 1 3

The data have been cbrrected for the constriction effects due to the presence of the tunnel w r i l l s . The corrections hkve not been modified to allow for the effect of SKeep. The FollEwEig €5E1e Shows €Fie G&iIb.de of the corrections tu Mach number ahd to dynamic pressure:

Corrected Mach number

0.210 .600 .80o .850 . goo 930 940

Uncorrected Mach. numb-

a. 210 .6oa 798

.848 896

=923 939

Tares

A correction to the drag data was necessary to allow for forces on the exposed surface of-tbe turntable. This cokke-ctfon was determined from tests with the model removed from the tui%tizble. The correction was found to vary with Reynalds in2iiber oxily a d is presented in the following table :

RXlO* C D ta2e

1.00 O.OO71 2 .oo .m63 3.00 ,0060 7.00 .om8 11.00 .0056

No attempt was made to evaluate tares due to possible interference effects between the model ahd the turntable.

TESTS

Reynolds Nruiiber Effect8

To determine the effects of Reynolds number on the aerodynamic characteristics of the horfzontal tail, lift, drag, pitching moment, and elevator hinge moment were measured for a Mach number of 0.21 at Raynolds numbers of 2,090,000, 3,OOO,OOO, 7 , ~ 0 , 0 0 0 , and 11,000,OOO. For these tests, the mgle-of-attack range vas from -LO0 to 24O, the elevator deflections were Oo, -loo, and -20°, and the tab was undeflected. For Mach numbers of 0.60, 0.80, and 0.90, similar data were obtained at Reynolds numbers of 1,000,000 and 2,000,000 with the elevator and the tab undeflected.

. .

.

UCA RM ~ 9 ~ 1 3 5

three separate sections. The seala were fitted closely to the ends of each section t o reduce leakage to a minlrmrm. The elevator was equipped with an unsealed tab, the Etrea of which was 6.5 percent of the elevator area and which extended fra 23.7 t o 44.8 percent of the semispan. The elevator and the tab gaps are shown in figure 1.

The stabilizer m8 constructed of solid steel and the elevator of aluminum alloy. The model was mounted vertically with the win&tunnel floor serving as a reflection plane as shown in figure 2. The rotating turntable upon which the mddel was mounted is directly connected t o the forc-easuring apparatus. The elevator and tab hinge moments were measured with resistasc-type electric straFn gages. The elevator gage was beneath the turntable coyer plates, and the tab gage was contained within the elevator. The elevator deflectian was remotely controlled and the tab deflection was set by means of an indexing system bu i l t in to the tab asd elevator. The gap between the elevator and the reflection p k e was approximately 0.02 inch when the elevator undeflected.

A streamwise row of orifices m s provided st 50 percent of the semispan t o measure the chordwiae distribution of stat ic pressure. S i x orifices were located in the balance chauiber, one on either aide of the B e a l a t 16, 4.8, and 90 percent of the semisp&n, t o measure the pressure differences across the elevatomose seal.

The data have been corrected for the effects of t u n n e l d i n t e r ference, for constriction due t o the presence of the tunnel walls, and for model-support tare forces.

T u n n e l - - W a l l Interference

Corrections t o the data for the effeots of tunnel-wall interference have been evaluated by the methods of reference 5 , using the theoretical span loading c a l c e t e d by the method6 of reference 6 . The corrections added. t o the drag and to the angle of attack were:

Aa = 0.329 CL, degrees

= 0.00502 m= No a t t q t was made t o separate the tunnel--wall interference effects

resulting from l i f t due t o elevator deflection and l i f t due to angle of attack. No corrections were applied t o the pitchingynaient or h u e - moment data.

6

Constriction Effects

NACA RM ~ ( 3 ~ 1 3

The data have been corrected for the constriction effects due to the presence of the tunnelwalls. The corrections have not been modified to allow for the effect of sweep. The following table show6 the magnitude of the corrections to Mach n&er and to dynamic pressure:

Corrected Uncorrected q uncorrected Mach number Mach m e r a corrected

0.210 0.210 .600 ,600 .800 9 798 . e o .848 .goo .896 - 930 923 9 b 0 932

Tares

A correction to the drag data was the exposed surface of the turntable. from tests with the model removed from was found to vary with Reynolds number following table :

1.001 1.001 1.002 1.003 1.005 1.008 1 .oog

necessary to a l low for forces on This correction was determined the +xr.mtable. - The -correction anly and is presented in the

RXlO* CD tare

1.00 0.0071 2.00 .0063 3.00 .0060 7.00 -0058 11.00 .om6

No attempt was made to evaluate tares due to possible interference effects between the model and the turntable.

Reynolds Mumber Effects

To determine tbe effects of Reynolds number on the aerodynamic characteristics of the horizontal tail, lift, drag, pitching moment, and elevator hinge moment were measured for a Mach number of 0.21 at Rq-nolds numbers of 2,000,009, 3,OOO,O, 7,00U,CSCjO, and ll,OOO,OOO. For these tests, the asgl+of-attack range was from -10' to 24O, the elevator deflections were Oo, -loo, and -20°, and the tab was undeflected. For Mach numbers of 0.60, 0.80, and 0.90, similar d a h were obtained at Reynolds numbers of 1,OOOjOOO and 2,0~,000 Kith the elevator and the tab undeflected.

.

NACA RM ~ 9 ~ 1 3

Mach Humber -Effects

7

.

To determine the effects of compressibility on the aerodynamic characteristics of the horizmtal tail, lift, drag, pitching moment, elevator hinge moment, tab hinge moment, pressure difference across the elevator-nose seal, and stremise distribution of static pressyre were measured at a Reynolds number of 2,000,000 at Mach numbers of 0.21, 0.60, 0.80, 0.85, O.gO, 0.93, an& 0.94. At Mach numbers less than 0.80, the angle"of-attack range was from -loo to 24O, and the elemto?+

' deflection range was fr& -25' to 6O. At Mach numbers above 0.80, the angular ranges were limited by nind-tramel power. Lift and hing- .. moment measurements were made with tab deflections of Oo, 5O, loo, and 13' throughout the cnmplete range of Mach nunibers and elevator deflections at uncorrected angles of attack of Oo, 4O, and 8'.

Effects of Standard Roughness and of Removal of the ElevatoHose Seal

Tests were a lso made to evaluate the sepgrate effects of standard. leading-edge roughness (reference 7), and. of removing the elevatomose seal on the lif't, drag, pitching+nment, and elevator hing-oment characteristics. D a t a were obtained at a Reynolds number of 2,000,000 over the angle-of'ttack range for elemtor deflections of kO, Oo, and -15O with the tab mdeflected at all test Mach numbers up to 0.93.

RESULTS AND DISCUSSIOH

The results of teste conducted to evaluate the effects of Reynolds number on the aerodynamfc characteristics of the horizontal tail are presented figyes 3 and 4, and results of tests conduct'ed to evaluate the effects of Mach number are presented in figures 5 through 12. The data from tests conducted to evaluate the separate effects of leadine edge roughness an& of removal of the elevatoIcnose seal are camred xith those obtained with the model in the norms1 condition in figures 13 through 16. An Index of the figures presenting the results is given in the appen-.

CertaFn data are presented for values of uncorrected angle of attack where:

a = 0.99 + Lm

!The constant 0.99 is the ratio between the geometric angle of attack and the uncorrected -le of attack indicated by the angle-ofkttack counter. The uncorrected angle of attack does not differ from the corrected value by more than 0.26~ for asy of the test points presented.

8 NACA RM A913

Effect of Reynolds Number .

Low speed.- l?he effect-s- of increasing the-Rejikilds number fr& " -

.. - .. 2,000,000 to 11,000,000 at a Mach nunber of 0.21 on the lift, drag, pitchinepmoment, and elevator hing-at characteristics are presented, in figure 3. The angle of attack at which the lift curve8 b e m e nonlinear was increased by approx-imately 3' when the Reynolds nuniber was increased from 2,000,000 to 11,000,000. This effect may be correlated with that sham in thepitching-mcrment data of figure 3(c) where the abrupt forward movement of the aeroaynamic center, which i e associated W L t h a loss of lift over the outer sections of a swept-back lifting aurface (reference 81, occurred at higher Uft coefficients as the Reynolds nuniber was increased. This delay in the loss of lift over the outer section was accompnied by more positive elevator hinge moment8 (fig. 3(b)), and by reductfans in the drag (fig. 3(d)) at the higher Reynolds nuuibers .

Increaaing the Reynolds rrumber caused very little change in the location of the .aerodynamic center (at C L ~ ) for elevator deflections of Oo and -loo. However, a forward movement of the aerodynamic center of 7 percent of the mean aerodynamic chord accompanied an increa13e.fn Reynolds number from 2,000,000 to 7,OOO,OOO with the elevator deflected -20'.

The close agreement of the lif't and elevator hingmoment coefficients obtained at the various Reynolds numbers over the linear range of the data of figures 3(a) d 3(b) indicates that C k , CQ~, C b a y and

were not sensitive to changes in Reynolds numbers between

2,000,000 asd 11,000,000.

The slope parameters measured from the results of tests of a gemlet- rically similar m o d e l conducted in the Ames 7- by 10-foot wind tumel (reference 2) are presented for ccmrparison with those evaluated from results of the present tests in the followJng table:

Slope Ames 7- by lckfoot parameter KLzaa tunnel

( refemce 2)

cLa 0.061 c%3 .032

--.OO24 -. 0078

Ames -foot pressure wind tunnel

0.059 .032 -. 0025 -. 0080

27.6 " . .. . . ..

A l l measurements of slope parameters were made From data. obtained at a Reynolds number of 3,OOO,OOO vith the exceptias of the values of

9 . CQ, and C& from the Ames 75LPoot pressure wlnd tunnel. These values

were measured from data obtained at a Reynolds number of 2,0003000. Be

The agreement between the lift and hing-oment parameters from the two investigations cazl be considered excellent. The reason for the difference of 2.6 percent of the mean aer-ic chord in the location of the aerodynamic center is not hm.

High subsonic SF&.- Figure 4 presents data obtained at Reynolds numbers of 2,000,000 and 1,000,000 at Mach numbers of 0.60, 0.80, and 0.90. These data show that the reduction of Reynolds number resulted in a reduction of lift-curve slope and a fo& movement of the aerodynamic center. The greatest effect occurred at a Mach number of 0.90 where the lift-curve slope was reduced by 0.003 per degree and the aerodynamic center at zero lift, was moved forward 2 percent of the m e a n aerodynamic chord. The change in Reynolds number from 2,000,000 to 1,OOO,OOO resulted in no important change in the drag for l i f t coefficients less than 0.4 or in the elevator hinge mament .

Effect of Mach Nunher

. The aerodynamic characteristics of the horFzantal tail at a Reynolds number of 2,000,000 for a range of Mach numbers f r o m 0.21to 0.94 are presented Fn figures 5 throqgh 12. - Lift.- The variation of lift coefficient with angle of attack is

presented i n . figure 5. At a Mach nmber of 0.21, the elevator wa8

effective in producing changes in lift throughout the elevatoAeflection and angl-f-attack range. As the Mach nunher was increased, the range of elevator deflections for which the elevator was effective at angles of attack greater than Eo was progressively reduced. It is not known if this effect exists at Mach nmbers great- than 0.85,as insufficient wind-tunnel power was available to test at angles of attack greater than 12’ at these speeds. The miation of lift coefficient with elevator deflection at an angle of attack of Oo is presented in figure 17. These data show that the elevator effectiveness was approximstely constant over a range of elevator deflectfons between *6O at all Mach number s .

The effects of Mach number on the values of CQ,, CL~, and, are shown In figure 18. The elevator-effectiveness parameter increased gradually f r o m 0.032 at a Mach number of 0.21 to 0.038 at a Mach number of 0.87; and decreased. abruptly with further increase in Mach number. The stabilizer effectiveness parameter C h increased from 0.059 per degree at a Mach n&er of 0.21to 0.082 per degree at a Mach number of 0.93 Fn close agreement with the variation predicted fran reference 6. Increasing the Mach number f r o m 0.93 to 0.94 resulted in an abrupt decrease of‘ C k A t low speeds, the value of %e was -0.54

e “%e

10 NACA RE4 Am13

and was little affected by compressibility at Mach numbers up to 0.70. At higher Mach numbers, the absolute value of % decreased, the decrease becoming very rapid at Mach numbers above 0.85. At a Mach number of 0.94, the magnLtude of %, had decreased t o 65 percent of its lmspeed due.

e

Hinge moment.- The elevator hinge-moment coefficients for various Mach nmibers up to 0.94 are presented in figure 6. as a function of angle of attack and in figure 7 as a function of elevator deflection. At the-higher Mach numbers the slopes of the curves vary considerably xith angle of attack and with elevator deflection; therefore, the hinge moment parameters Ch and C k are not indicative of the hinge+ moment characteristics of the horizontal tail and any discussion Fn terms of these parameters would be misleading. At Mach numbers less ..

than about 0.85, increasing the Mach number caused gradual changes in the elevator hing-ment coefficients for elevator deflections and angles of attack between 6'. The Mach numbers at which rapid changes in the elevator hinge-mment coefficients occurred were dependent upon the elevator deflection and-the angle of attack. This is illustrated in figure lg(a) which presents the variation of elevator hinge-moment coefficient with Mach number for several angles of attack at 0' elevator deflection and in figure lg(b) which presents the variation of elevator hing-oment coefficient with Mach number for several elevator deflections at an uncorrected angle of attack of Oo.

ese

Tab effectiveness.- The variation of elevator hinge-moment coefficient with elevator deflection for several tab deflections is presented Fn figure 7. The tab-effectiveness parameter , measured at OO elevator deflection, had a value of -0.0035 and was little affected by compressibility. This is evident from the data of figure 20 which presents the elevator hing-ment coefficient produced by tab deflection Ache as a function of Mach number. For negative elevator deflectlorn the tab effectiveness decreased with increasing Mach number, especially at the larger tab deflections. With a n elevator deflection of -loo the tab was ineffective when deflected more than 10' at Mach numbers above 0.90. The change in lift coefficient due to deflection of the tab is shown in figure 17.

. .

The tab hing-ment coefficients are presented in figure 8 to permit application of the tabeffecti'veness data to the design of a simple or. spr-tab installation.

Pressure difference acros~ the elevator-nose sed.- The v a r h t i o n with elevator deflection of the pressure coefficient across the elevator- noae seal is presented in figure 9. The low-speed data are not in close accord with those of reference 2. The comparatively greater spanwise variation of pressure coefficient across the elevato-ose seal than is shown In reference 2 may be attributed to the division of the balance chamber into three sections between which aifcould not flow. The hinge

..

acLcB RM Awl3 ll

line of the model tested in this investigation was slightly offset, which required a distortion of the curtains t o prevent a dLscontinuity at the hinge line. Evidence of %his &set of the elemtor hinge l ine fs given by the preesur&stribut€on data shown i n figure 12. Leakage around the ads of the seal in each section of the balance chamber may have had an effect on the balancing preseures, particularly tn the t i p section of the balance chauiber *re the rat io of leakage area t o vent area between the curtains and the elevator was the greatest.

Inspection of the data presented in flgure 9 reveals that, in general, the rate of c.$ange of the pressure coefficient across the elevato?xuase aeal with elevator deflectian, measured a t Oo elevator deflection, became more positive as the Mach nu&= was increased t o 0.93 . At a Mach nuniber of 0.21, the rate of rise of halancing pressure with elevator deflection decreased abruptly a t large negative elevator deflections. A t Oo angle of attack, for erample, the balancing pressure in the middle section of the c-er did not increase when the elevator was deflected m r e than -20a. As the Mach nWer m a increaaed, a decrease of balancing effectivenees occurred at progressively smaller elevator deflectians. At a.Mach mer of 0.93 and at- an angle of attack of Oo, the balancing pressure in the m i d d l e chaniber increased l i t t l e as the elevator was deflected beyond -4O.

* In order to evaluate the reduction of elevator hinge moment obtain- able through the use of a sealed internal aeroayneslic balance, the -&t coefficients for an elevator ~ i t h a balance p h t e extending' fram 0 t o 96 percent of the span with a chord of 0.35ce' have been computed. The total elevato3deflection rasge would be limited t o approximately go if' this amount of balance #ere employed. In computing the hing-t chsraderist ics of the balsnced elevator, it was assumed that the pressure differace indicated by each pair of orifices existed uniformly over the balance plate betrreen the center lines of the hinges which limited that sectian af the balance chamber wherein the orifices were located. The computed hinge mmenta of the balanced elevator m e compared xith the measured hinge mments of the radi-ose aealed elevator in figure 21. These ccaagutations ahox that, a t a Mach number of 0.21, use of the - sealed internal balance would result in a

As the Mach number was &&sed, the range of elevator def lectian f o r

pr0gressiGel.y decreased. A t a Mach mnder of 0.93, for example, the

value of - w&s reduced by only about 12 percent at elevator deflec-

tians greater than 4'. At this Mach rimer the value of - was

approximately Zen for elevator deflections between *lo. Any greater

ache b e

&he ase

12 HACA RM ~ 9 ~ 1 3

amount of internal balasce would result in overbalance for small elevator

deflections. The range of angle of attack for which was reduced by the incorporation of the Fnternal balance became progressively smaller a8 the Mach nuPiber was increased.

&he

Pitching moment.- The pitchiqpuaent coefficients about the quarter point of the mean aerodynamic chord are presented as a function of lift coefficient in figure 10. At low speed,loss of lift over the outer sections, which caused static longitudinal instability as indicated by the break in the pitch-merit curves, occurred at a lift coefficient of about 0.6 with tbe elevator undeflected and at lower lift coefficfents as the elevator was deflected negatively. A n increase Fn Mach number to 0.9 caused little change in the lift coefficient at which static longitudinal instability occurred. This instability did not occur at Mach numbers of 0.93 and 0.94 within the angle-of-attack range for which data were obtained.

The rate of change of pitching-mament coefficient with lift coefficient shows that the static longitudinal stability of the horizontal tail increased aa the Mach number was Increased. This effect is 8lllmn~rized In figure 22 where the location of the aerodynamic center (for Se=Oo at C L ~ ) is presented 88 a function of Mach number. At law speeds, the aerodynamic center was at 27.6 percent of the mean aerodynamic chord asd its location varied only slightly xfth Mach number up to 0.85. A rapid rearward movement of the aerodynamic center occurred as the Mach nmber was increased from 0.85 to 0.94.

The pitching moment due to deflection of the elevator increased with Mach nzmiber as illustrated in figure 22, where the negative value of the parameter C is shown to increaee frcm-O.OOg8 per degree at a Mach number of 0.21 to -0,0165 at a Mach number of 0.94.

%e

&a&.- The drag data of figure 11 are summarized in figure 23 where the minimum drag coefficient, maxfrmrm lift-drag ratio, and the lift coefficient at which the maximum lift-drag ratio occurred are presented as a functian of Mach nmiber. The Mach number for drag

divergence, defined as the Mach number at which = 0.10, was approximately 0.91 when the elevator X&E undeflected. The maximum lift-drag ratio was 20.5 at a Mach number of 0.21. The effects of ccmpressibility on the naximum lift-drag ratio were small up to a Mach number of about 0.80, but marked decreases occurred with further increases in Mach number. At a Mach number of 0.94 the maximum lif-ag ratio war3 about one-third of that at low speed. The lift coefficient for maximum lift- drag ratio increased with Mach number at Mach numbers greater than about 0.60.

ac,

NACA F34 Awl3 13

.

Pressure distribution.- The streamwise distribution of static pressure at the midsemispan was measured to correlate the effects of Mach number, as evaluated from force measurements, with changes in the surface pressures and to provide d a t a for structural desi-. These pressur- distribution data are presented Fn figure l2 for various elevator deflections and angles of attack for the same Mach nmibffs for which force data are presented.

The magdtude of the pressure coefficient which corresponds to sonic velocity normal to the quartexhord line was calculated from the following equation based on simple

where 7 is the ratio of specffic

Bweep theory:

L M2 cos2 3') - 1 ]

7+1

heats end is equal to 1.4. The values of Pcrn calculated from the above equation are shown Fn

figure 12 in order to indicate the condition8 for which there was supeF sonic f l o w in a direction normal to the quar-hod lFne at the

- 5O miasemisp&n.

The reasan for the reduction of elevator effectiveness at the higher Mach numbers is eddent from the data of figure 12. At low speeds, the decrement in lift due to negative elevator deflection was distributed over the airfoil chord. However, at the higher Mach numbers, deflectLon of the elevator produced very little change in the surface pressures fo-d of points on the stabilizer where the flaw was indicated to be supersonic normal to the quarte-ord line. At emall elevator deflections, increasing the Mach number beyond 0.9 caused supersonic flow over the elevator. The resultant change in the load distribution over the elevator can be correlated xfth the large Increase in elevator hing-t coefficient shown in figure lg(b) for an elevator deflection of -4' between the Mach numbers of 0 . 9 and 0.94.

Effects of "Edge Roughness

Results of tests conducted with standard roughness, applied to the leading edge as described in reference 7, are presented In figures 13 through 16. Results of tests without lea-dge roughness are d s o presented in these figures for purposes-of comparison.

LOSS of lift over the outer sections of the tail occurred at slightly lower angles of attack at Mach numbers greater than 0.60 when leadiwdge roughness was applied. (See figs. 13 and 15.) The stabilizer effectiveness and elevator effectiveness were reduced when

roughness wae applied t o the Leading edge. These effects axe summarized in figure 24 ,where the values of C h and Cue, xhich were measured f r o m the data of figure 13, axe presented as a function of Mach nmber . It was asaumed In measuring C%, that the elevator effectiveness was constant between deflections of 0' and 4O. The greatest reduction in effectiveness occurred at a Mach number of 0.93 where C r , and Cue were each reduced by 0.012 per degree.

Inspection of the data of figure 14 shows that leadiwdge raughness cauaed sizable reductions in elevator hinge moments when the elevator vas deflected.

Leading-edge roughness caused a large forward shift of the aerodynamic center at Mach rimers greater than 0.70. Figure 25 shows thie effect to be greatest at a Mach nmber of 0.93 where the aerodynamic center (measured for 6& at CL-) was shifted fram 36.7 to 28.8 percent of the mea aerodynamic chord.

AB would be expected, application of leading-edge roughneea resulted in increaaed drag. (See fig. 16.) Figure 25 shows that the increment of minimum drag coefficient due to leading-edge roughness was about 0.0060 at low speed and about 0.0040 at a Mach n W e r of 0.93.

Effect of Removal of ElevatoIcElose Seal

The effect of remarral of the elemto-ose aeal fa shown in figures 13 through 15 where conparimm is made between data obtained with the elevator noBe sesled and with the elevator nose unsealed.

The data of figure 13 show that unsealing the elevator nose had no important effect r m the variation of lift with asgle of attack. The elevator effectiveness, however, wa8 reduced. This effect ie summarized in figure 24 where CQ, is presented as a functian of Mach number. The maximum reduction in C Q ~ was 0.004 per degree, which occurred at a Mach number of 0.60.

IR general, unsealing the elevatol.-nose gap caufied alight increases in the elevator hinge mament when the elevator was deflected.

The results of KLnd"tUnne1 tests conducted to evaluate the independent effecte of Reynolds nmber and of Mach number on the aerodynamic characteristics of a horizontal tail of aspect ratio 4.5 +!:h the quartel.- chord line swept back 35O have been presented.

HACA RM ~ 9 ~ 1 3 1.5

.

Results of testa at a Mach number of 0.21 over a ranQe of Reynolds numbers f r o m 2,000,000 to ~,OOO,OOO tndicated thatr

1. An Increase of Reynolds number increased the asglc+of+ttack range mer which the variation of lift Kfth angle of attack was linear.

2. The aer-c characteristics of the horizontal tail were not sensitive to scale between Reynolds n'Lmibers of 2,000,000 and U,OOO,OOO within the angl-f-ttack range for which the miation of lift with angle of attack was line-.

Results of tests at a Reynolds nmiber of 2,000,000 over a range of Mach nmbers f r o m 0.21 to 0.94 3ndicated that:

1. An Increase inMach m e r frcw 0.21 to 0.93 result& in an increase in lift-curve skqe f r a m 0 -059 to 0.082 per degree; further increase in Mach number to 0.94 caused as abrupt decrease -In lifkurve slope.

2. The elevator-effectiveness parameter CQ, increased. fram 0.032 to 0.038 per degree between Mach nmibers of 0.21 and 0.87, and diecreased rapidly as the Mach number was incream?d to 0.94.

3. The Mach nunher at xhlch marked chasges in the elevator hinge- m'ent coefficients occurred was depedent upan the angle of attack and elevator deflection; hawever, the changes in the elevator hinge-moment coefficients at angles of attack and elevator deflections between 6' were gradual a8 the Mach nunher uaa increased to 0.85.

4. The tab was effective in productng a balancing increment in elevator hinge mament throughmt the Mach number range.

5. Incorporation of sealed internal balance sufficient to cause a 5+percent reduction in the variation of elev&tor hinge moment wlth elevator deflection et a Mach rider of 0.21 caused only a -percent reduction at 8 Mach nmiber of 0.93 for elevator deflections greater than 40.

.Results of tests made to evaluate the effect of LemWneedge roughness indicated that:

1. L e a w d g e roughness caused reducticma in the l i f h m e slope and in the elevator effectiveness.

2. L e a m e roughness caused a eizable reductian in the elevator hinge mment when the elevator was deflected.

Results of tests made to evaluate the effect of unsealing the elemto-ose gap indicated that:

16 HACA RM Awl3

1. UnSealLng the elenatmoas gap had no tqortant effect on the 1if"e slope, but reduced the elevator effeotivenees.

2, In general, unsealing the elevrttor-noae gap caused alight increases in the elevator hinge llLQment when the elevator m e deflected.

Ames AerQnautical Laboratory, Hati& Advisory CoPnnittee for Aeronautice,

Moffett Field, Calli.

The following tables have been included to provide a convenient index to the data of this report:

FORCE AND MCMEKT CHARACTERISTICS

Reynolds Number Variable

Results presented R

2,000,000

l l , O O O , 000 to

1,000, OOO d

2,000,000

I

M

0.21 I 0.60,0.80,

0.90

-10 t o 24

V

0,-lO,-QO

1 0

1

6 t t

0

t

WCA RM ~ 9 1 3 Mach Nuiber Variable

M 0 -21

-60 -80 85 -90 93 94 .2l

-60 .80 -85 90 93 - 94

-21 0 6 0 .80 85 90 93 94 21

a 6 0

-80 a 8 5

0 9 0

93 94 .21 -60 80 85 .go 93 94 21

I -60 -80 -85 -90 93

9 94 .21 -60 .80 S 83 -90 93 94

=

[R=2,0C

a, aeg

-10 t o 24 -

\1 -10 t o -20 -10 to I 2 -10 t o 10 - 0 t o 8 -10 to 24

3. -10 t o 20 -10 t o 12 -10 t o 10 - 8 t o 8 0,4,8

1 0 4

o b .1

014, 8

1 -8 to 24

.1 to X)

4 t o 8 - 8 t o 8

-10 t o 24 At0 a

\L -10 to 20 -10 to 12 -10 to 10 - 8 t o 8 -10 t o 24

4 -10 to 20 -10 to 12 -10 to 10 *to 8

two 1 Ee, *g

6 to -25 -

.1

1

1

6 t o -20 6 to 40 4 to -10 6 t o +

6 to -20 6 to -20 4 to -10 6 t o - a

6 to40 6 to -20 4 to "10 6 to -25

5. 6 to -25

.1

.1

.1

1

6 to -25

6 t o -20 6 to 40 4 to -lo 6 to -25

6 to -20 6 to "-20 4 to -10 6 to -25

6 to -20 6 to -20 4 to -10

0

Y 0 to 15

V 0

V

Figure

18 NACA RM ~ 9 ~ 1 3

STREAMWISE DISTRIBUTIOm OF STATIC PRESSURE AT TEE MIDSEMISPAN

[R=2,000,000; 6t=0° f

Results presented

P YS percent chord

- M

0.21

.21

.60

.60

.80

.80

85

85

a 9 0

93

94

-

-

12,16,20

0,4,8

12,16,20

0,4,8

0,-4,-10 -15 ,-20

0,4,-10

NACA RM ~ 9 ~ 1 3

c

"10 to 24

1

\1

1

.1

-10 to 20 "10 to 12 -lo to 10 -10 to 24

-lo to 20 -lo to I 2 -10 to 10 -10 to 24

"10 to 20 -10 to 12 -m to 10 "10 to 24

-10 to 20 -10 to 12 -10 to 10

.

20 RACA RM A9313

Results presented a, deg - 0

""_ -8 to 24

0

0

-10 to 24

""-

""- ""_ ""_

""

0

6 to -25

0,-6,-10

6 to

""

o to -15

""

""

o to 15 ""

0

0

5,10,15

0

""

0

""

?Shows the computed effect of a sealed internal aerodynamic balance on the elevator hinge-mment coefficients.

2 Shows separate effects of 1eadiIqpedge roughness and removal of . elevator-nose seal .

HACA RM Awl3

1.

2.

3.

4.

5.

6 .

7.

8.

Do*, W e s B., Jr.: WInd4bmel Investigation of Horizontal !Tails. I - Unswept and 35' Swept-Ekck Plan Forms of Aspect Ratio 3. HACA RM A7K24, 1948.

Dods, Jules B., Jr . : WInd4Tmnel Investigation of Horizontal Tails. I1 - Urnwept and Bo SwepGBack Plas Forme of Aspect Ratio 4.5. aACA RM A m , 1948.

Dods, Jules B., Jr.; Winddmnel lbvestigatian of Horizontal Tails. 111 - Unswept and 35O SKept4Jac.k Plan Form6 of aspect Ratio 6 . NACA RM A8E30, 1948.

Doh, Jules B., Jr.: W i n d a e l Investigation of Horizontal Tails. N - Umwegt Plan Form of Aspect Ratio 2 and a Tuc+ Dimensional Model. 'HACA RM A w l , 1948.

Sivelh, James C . , and Deters, Oven J. : JetGf3omdary and Pl81+ Form Correctians for PartialSpan Models With Reflectfon Plane, End Plate, or 1Qo End Plate in a Closed Circular Wind !Pmnel. HllCA Rep. 843, 1946.

DeYoung, John: Theoretical Additianal Span Loading Characteristics of U h g s With Arbitrary Sweep, Aspect Ratio, and Taper Ratio. XACA TIT 1491, 1947.

Abbott, Ira H., von Doenhoff, Albert E., and Stivers, Louis S., Jr.: smumary of Airfoil Data. . RACA Rep. 824, 1945.

Shortd, Joseph A., and Maggin, Bernard: Effect of Sweepback and Aspect Ratio on Langltudlnal Stability Characteristics of W i n g s at Low Speeds. NACA M 1093, 1946.

22 NACA RM ~ 9 ~ 1 3

Upper and Lower Surfaces

Statim, Ordinate

.50 0 7 5

1.25 2.50 5-00 7.50 10 00

20.00

30.00

15.00

25.00

35 00 40.00 45.00 50.00 55 00 60.00 65.00 70.00

80.00 83.00 go. 00 95.00 100.00

75 00-

L. E. radius, T.E. radius,

0 .8Q4 969

1.225 1.688 2 327 2.805 3.199 3 813 4.272 4.606 4 837 4.968 4.993 4.894 4.684 4.388 4.021 3 9 597 3.127 2.623 2.103 1.582 1.062 .541 .021

.

m I

Diinemions shm, in inches unless otherwise noted

. . . . . . . . . . . . . . . . . . . . . . . . .. . . - . . . . . . . . . . . . .

, ,

Aspect rafio

025 chord of airfoil section Toper ratio

Area semispan Elevator hinge line, chord of airfoil secfion

Elevator area

Tab ar8a

Locatim of pressure- dtstdwfim oriffices

NACA 64AOlO airfoil secfion

F

4.5 s 0.5 G

4.443 ffP

1.204 ff

0.0777 ftp

1,458 ft

-25.7 Secfion A - A

. . .

NACA m ~ 9 ~ 1 3 25

Figure 2.- Semispas horizontal tail model mounted in the %foot pressure wind. tunnel.

W A RM Awl3 27

.6

.4

- .6

- .8

9.0

-/2 -8 -4 0 4 8 /2 /6 20 24 28 Angle of affack, Q I deg

v

Figure 3.- The effect of Reyno/ds number on fhe /ow-speed aerodynumic characteristics. hf, 0.24 S,, 0:

2a HACA m . ~ g ~ 1 3

. LO

.8

.6

.4

* .2

-.2 4

-.4

-.6

-. 8

-L 0

-/. 2 .20 . ./6 . /2 .08 .04 0 -. 04

Pif ching-momen f coefficienf , Cm - IC) Gr vs Cm .

Figure 3- Confinued.

. . . . . . . . . . . . . . . - . . . .. . . . "

I

. . . . . .

. .. . . ' ' ' I .

!

I

RACA RM ~ 9 ~ 1 3 *

NACA RM ~ 9 ~ 1 3 *

1.0

.8

-6

-4

-2

0

-. 2

-.4

-. 6

-. 8 .04 I

33

i

t

1 i c

i -.04 508 for M= 0.60 -qggi7 Pifching-mment coefficient , C ,

( c / c, vs cm . Figure 4.- Continued.

..

0 .04 .08 ./2 .I6 .PO 9 4 .28 .32 .36 .# .44 for M= 0.60 Drag coefficient, CD

. . . . . . . . . . . . . . ... . ! ' . . . . . . . . . . . ... . . . .

NACA RM AgG13 35

fa) M, 0.2/.

NACA RM A9213

(&/ M, 0.60.

NACA RM AgG13

Figure 5.- Confhued.

37

(c) M, 0.80.

38 NACA RM ~9313

-/2 -8 -4 0 4 8 /2 /6 20 24 Ang/e of of fock, 0, deg

(d) M, 0.85.

f igwe 5. - Continued.

.

NACA RM A9313

/. 0

.8

.6

.4

0: .2

s o

% C

0 -2

Q 8 > -.2 4

4

-. 4

-. 6

-. 8

-/. 0

-I. 2 - /2 -8 -4 0 4 8 /2

39

- Ang/e of affuck , a, deg =Ez7 /6

- l e ) M, 0.90.

Figure 5.- Cunt inued.

NACA RM ~ 9 ~ 1 3

I. 0

.8

.6

.4

(r

-2

.2 x' 0

8 0 5 8 3 -*2

-.4

-.6

-. 8

-LO

-1.2 32 -8 -4 0 4 8

Angle of atfock, 4; deg

T g g 7 ( f ) M8 0.93.

Figure 5.- Cont/irued.

NACA RM ~9313

.8

.6

.4

.2 e z 9 0

g o 5 -.2 8

"4

-.6

-.8

41

-8 -4 0 4 8 /2 Angle of of fock, U, deg .

I g ) MI 0.94. =-557 - Figure 5.- Conc/uded.

42 HACA RM AgG13

-/2 -8 -4 0 4 8 /2 16 20 24 28 Angle of attack, Q , deg

" (b) M, 0.60.

43

.40

.32

NACA RM ~ 9 ~ 1 3

(c) M, 0.80. - figure 6.- Confinued.

'2 -8 -4 0 4 8 12 16 20 24 28 Angle of oftock , Q I deg

HACA RM AgG13 45

.44

.40

.36

-32

.28

8 -24 5

-12 -8 -4 0 4 8 /2 /6 20 24 Angle of uf fuck , P , deg

(dl M, 0.85.

Figure 6.- Continued.

4 6 - NACA RM A9G13

Y

-48

-44

.40

.36

-32

-5 .04 c

42 -8 -4 0 4 8 /2 Angle of offuck, Q , deg

42 -8 -4 0 4 8 /2 Ang/e of of t a c k , (I, deg

( f ) MI 0.93. ( g j M, 0.94.

Figure 6.- Conchded.

. . . . . . . . . . . . . . . . . . . . - . . . . . . . . . . . . . . .

Efevafof deflecfim, 8, deg

l a ) At, 0 , P f . Figure Z- The variafion of ekvafor hinge-moment coeffcienf wlfh efevafor Mfecfion. R, 2JW@0.

. . . . . . . . . . .

i '

. .

. . . .. . . . . . . . . . .. - . . . . . . . . . . . . . . . . . . . . . . . . . . ' I

I I

.28

-16 -28 -24 -20 -i6 42 -8 4 0 4 8 for ~r, = Oo

Elevator deflection, 8#, dog -a37

/b ) M, 0.60. F&um Z- Continued

P

4 P w

UI .F

..

t

48 -24 -20 46 -12 -8 -4 0 4 8 for a, = 0 0

Elevotor deffecnon, S,, deg v (d) 111, 0.85.

F&um Z- Continued

. . . ..

. . . . .

. . . . . . . . .. . . . . . . . . .

-24 -20 -I6 42 -8 -4 0 4 8 for Q, = 0 0

Elevator deflection, 8, , deg

{el M, 0.90. v

F&um T- Continued.

. . . .

. ..

..". .

NACA RM ~ 9 ~ 1 3

4

I

.36

-32

28

-24

.20

. /6

. /2

.08

.04

0

7 0 4

-.08

: /2

: /6

72 0

724

728

-.32

53

-20 -16 -/2 -8 -4 0 4 8 for ~ r , = Oo

Hevutor def/ecfion, 8, , deg T ( f ) M, 0.93.

54

.20

. / 6

8 * I2

$ s. C .-2 .08

Qr 8 .04 *z s e o e & -s

2 -.08

0

-.04 -e L

e G Q

-. /2

-. 1 6

-20 -12 -8 -4 0 4 B

Elevutor def/ection, Se ,

/g/ M, 0.94.

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

"

I I I I I I I I ~ t i i i i i i i i -2420-16424 -4 0 4 -2420-16-12 -8 -4 0 4

Etevafor cbflectrbn, a,, deg

(a) M, 0.21 to M, 0.80,

FQum 8- The w&Mn of tab hlngwmnwnt mfflcient with elevator &flecffon. R, 2,000,OUO.

. . . . . . . . . . . . .

.24

20 .I6

.I2

.08

.04

.. . . . . . . .

1

-24-20-16-12 -8 -4 0 4

& I deg

(b) M, 0.85 to #, 0.94.

-12 -8 -4 0 4

-5v

. .. . . . . . . . . . . . . . . . . .. . . . . . . . . . . . . . . . . . . . . . . . . .

1

. . . . . . . .

L

B B

w

. .

.. . . . . .

€levo?br deflecfion , &, deg

( b ) M, 0.60.

Fgm 9.- Continued.

. . . . . . . . . . . . .. . . '!

.

Figure 9.- Confinued.

. . . "

1 I

- .6 \ p .4

2 -.8 8 40

e cr, Y 9

8 -2

4 \ t3 e w e C I

8

P F

ir,

(e) M, 0.90. Figure 9.- mnfinued.

Elevator deflection, L%, deg

(fl Ad;. 0.93.

Figure 9.- Concluded.

-12% -4 0 ‘ 4 -12-8 -4 0 4

Elevator deflection, &, deg

(91 M, 0.94

. . . .. .

NACA RM ~ 9 ~ 1 3 63

LO

-8

.6

.4

0:' .2 4 C

a?

$ 0 0 Q

"2 4

-.4

-. 6

-. 8

-/. 0

-1 2 .24 .2U ./6 . /2 . 0 8 .m 0 -.04 -.08

t u ) M, 0.2/.

Figure /O.- The variufion of /iff coefficient with pitching-moment coefficient. 81, Oo; R, 2,000,000.

64 " . . . . - NACA RM ~ 9 ~ 1 3

.8

.6

.4

.2 %* C -9 s o

5 e 8 % 0.2

-.4

- .6

- .8

-10

-I 3 "f20 ./6 ./2 .08 . 04 0 -. 08

.. (e/ A4, 0.80.

Figure 10.- Confinued.

66 I - NACA RM ~ 9 ~ 1 3

(dl M, 0.85.

Figure 10.- Confhued.

LO

.8

-6

.4

Q4

-9 .% 4 .2 e

$ 0

5 0 QI

CJ

.2

- .4

- .6

- .8

-LO

%2 .24 20 ./6 ./2 .08 -04 0 704 708 :/2 :/6

Pitching-momenf . coefficient, C' =igzJ7

- (e/ M, 0.90.

- Figure 10.- Confinued.

.. . . . . . . . . . . . . . . , . .. . . . . . .

LO

.8

.6

.4

'312 z d C

$ 0

8 k B

5 -2 - .4

-.6

- .8

i . .

32 .28 2 4 20 .I6 .I2 -08 .04 0 :04 ~ 0 0 :I2 716 -20 2 4 Rtmllirg-mananf t w e f k q c;r,

TE&7 ( f ) Ai, 0.93.

Figwe 10.- ConHnued.

I i 1 : . . . . . . . .

NACA RM Awl3

.8

.6

.4

.

- -4

- .6

- .8

-LO .24 .20 .I6 .I2 .08 .04 0 -04 708 -12 716

pifching-momenf coeHicim4 C;, (g) M, 0.94.

Figure /O.- Concluded.

0 134 .08 .I2 .I6 10 2 4 28 .32 .36 .40 .44 Drag co&ficie& Go

( a ) M, 021.

Figure 11.- The . . vof/ot/on of /Iff coefficient with drag coefficient. 8t, Oo; R, 2,000,000.

8 I

. .

4 0

. . - - . . . . . . . .

I 1

". . . . . . . . . . . . . . . _ - . . . .

1

.8

6

.4

et * .P

* -2 3

-.4

-.6

-.8

I D4 R8 .I2 16 40 24 28 -32 .36 .40 .44 Drug mdficiient, C,

/ c ) M, 0.80. yu7

Figure 11.-Continued.

.. . . .

I I

. . . . . . . . . . . - . . . . . . ..

I

.. .

-t. .8

6

-.4

-.6

-. 8

-LO

# .08 ./2 J6 .20 24 28 .32 0 20 24

Figure It,- Confinued.

I . . . . . . .. .

I

... .

. . . . . . . .

I I

. .. .

. . -

I

4 P !!

-&? Y)

-a -6 -4 -2 0 d .4 .U #

4-40 &--I9

W i i i i i i i i i l

. . . . . . . . . . . . . .

, . .. .

I

. . . . . . . . . . . . . . . . .. . . ..

I L

- . . . . . . .

I I

. .

. .

&/O'

Q

B h

&-4.

0 m l O d D w

i

i !

4"'

F i i i i i i t T l I

4

F F

I

. . . . . . . - .

I

. . . . . . . . . . ... . -

, . . . . . . . .. . .

@ #

. ..

&=0'

Rgure 12.- CoMnued.

I

. ..

8'-4'

o p o 4 o M ) d o

8*=-100

1 I ..

. . . . . . . . . . . . . . . . . . . . . . . . . .

I

. . . . I

-12 -uI -8 -J5 -.4 -2 0

.I I

.6 d

o m u ) b p m

w

4

Flgure 12.- ConMwed.

.. . . . . ..

Figure E.- Continued

Q

S*=O" -12 -lo -.8 -8 -.4 -2 0 2 .I .6 B

-Lp

-ID -s -.!3 -.I -2 0 P .I .6 B

-12 -m -8 -6 :I -2 0 P .I 6 E o P o 4 o m m

8,"4'

0 p O 4 0 6 0 8 0 0 Po 40 60 80

8, =-20"

F n

Figure 42.- Continued

. . .

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

I I

8, = 0" -lo -B -8 -.4 .-2 0 P .4 .6 B

. . . . . . . . . . . . . . . . . . . . . . . . . . . . -

I I

&-$"

figure 12.- Concfuded,

. . . . . . . . . . . . . . .

6 "I

. .

-12 -8 -4 ~ 0 4 8 12 M 20 W for Be.= Oo s Angle of attuck, U, deg

(a) Af, 0.21.

\. c 52

I

Angle of attack, a, deg

(b) M, 0.60. v

Fgure /3.-*Cont/nued

. . . . . . . . . .. . . .. . . .

I

I

I L I . I

. . .. . . . . . . . . . . . . . . . . . . . . .. . . . . . . . . . . . . . . . . . - - . . . . - . . . . .. . . . . I

L I

f & w e A- cantinued.

. . . . .. 1

. . . . . . . . . .. . . . . . . . . . .

.8

.6

.4

-.4

- .6 - .8

42 -8 -4 0 4 8 lP 16 for Be= Oo

Angle of attack, a, deg

/e) At, 0.90. - 5 w

. 1 . . . . . . . . . . . . . .. . . . . . . . . . .

.- . . . . ... ."

! # e

.8

.6

.4

.2

0

- .2

- .4

- .6

i 1

. . . . . . . . . . . . i

-12 -8 -4 0 4 8 12 for Be= 0 0

Angfe of oltock, a, deg

( f ) M, 0.93. v

. .

92 NACA RM ~ 9 ~ 1 3

I

24

20

. 16

. /2

.08

.04

0

-.04

708

-20

724

-12 -8 -4 0 4 8 12 16 20 24 28 Angle of at tack , CI, deg

(u) M, 0.2/. v Figure /4.- The independent effects of /euding-edge rough-

ness und remowd of the elevutor-nose seal on the variation of e/evutor hinge-moment coefficient with mgVe of uffock. t+# og R, 2#000,000. *

NACA RM As13 93

-24

20

./6 8

4

-12 -8 -4 0 4 8 12 16 20 24 28 Angle of aftock # 4 , deg

I

=E27 R (6) M, 0.60.


28

2 4

20

0 . /6 s

-28

. - . . . . .

-.32 -/2 -8 -4 0 4 8 /2 /6 20 24 28

Angle of ottock Q deg

=Kjg&== .. "

(c/ M, 0.80. r


NACA m ~ 9 ~ 1 3 95

Angle of uffack, CI, deg

v (d l M , 0.85. figure 14.- Continued.

NACA RM ~ 9 ~ 1 3 I.

.


-/2 -8 -4 0 4 8 /2 16 20 Angle of attack, P # deg

( f ) M, 0.93.

97

figure /4.- Conc/iudea!

l0

.8

.6

.4

Ui 4 * .2 -? e .Q

QI 0 tr

$ 0

3 -k -.2

- .4

- .6

- .8

-1 0

L. E rough, seo/ed

.20 ./6 ./2 .08 .04 0 0.04 for = 0 O

PiMing-mmmt coefficien4 &

(a) MI 0.3/.

Figure /5.- The independent effects of /euding-edge roughness und remmul of the e/svcrtor-nose sea/ on the vuriufion of lift coef-

ficient with piichng-moment caffiiient. 0'; Rp ~00~000. a

NACA RM A9313 - 99

-

.

-20 ./6 ./2 .08 .04 0 0.04 for * 00 Pitching-mmenf coeMck4 C;,

f b ) M, 0.60. "

Figure 15.- Continued

100 MACA RM AgG13

.

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -

I b

. . .

. .. . . . . . . . -. . . . . . . . . . . . . . -

.. . .

~. . . . . . . . . . . . . . . . . . . . . - . . . .

I I I'

.8

.6

.4

- .4

-.6

- .8

. . . . .. . . . . . . . -.

, L

28 .24 .20 J6 .I2 .08 .04 0 4 4 4 8 -.I2 for Se = Ob mQ"t t.vem%w{ G

( f ) M, 0.93. [email protected]

Figure 15-Concluded. P W 0

. , " . .

LO

.8

.6

- .4

- .6

- .8

i,O 0 .04 .08 .I2 .I6 .20 .24 .28 .32 for ad = 0' -

Drug coefficie'ent, Go

(a) M, 0.21. v

c I

LO

.8

.6

.4

- .9

- .6

- .8

-LO

. . . . . . . . . . . . .

. ...

LO

.8

.6

.4

Cjj

*? G O a ‘4 -2 h 4.

‘4 . .2

C

c,

k PI

c,

- .4

- .6

- .8

-10

Drag coefficien4 C,

figure 16- Continusd.

*

- -. . . . - . . . . . . . . . .

L .. .. . . .

I

. .

r

- " . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -. - . . . . . . . .

c

LO

.8

.6

.4

" .4

- .6

- .8

Figure 16- Continued.

. . . . . .

. . . . .

I t

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . - . . . . . -. . . . . . .. . . . . . . . . . . .

0 .04 .08 .fP ' ./6 for 0' 0 .04 .08 ./P .I6 for' %,- 00 Drug coeffichnt, C, Dmg coefficien4 C,

. . . . . . . .

a

. ..

. . . . . . . .

NACA RM "

.4

.2

0

-. 2

Q* -.4

-28 -24 -20 -/6 -/2 -8 -4 0 4 8 E/evufor de f/ec fion, Se, deg

4

Figure K- The worhfion of /iff coefficient wifh e/ewafor def/ecfion. . QU, 0; 6 2,000,000.

110 NACA RM A w l 3

. . .

RACA RM A s 1 3 UI

/6

./2

Q,+Q .08

%* .04 42 Q

g o e 0 CJ -.04 % C Q

e -.08 B I Q) -.I2 -s

-. /6 8 P G

-e

%

p, -10

-24

-28

-. 32

QU 'deg -8 -6 -4 -2 0 2

I l t I l I I I o 4

A V D

V 7

a

/ .2 .3 .4 .5 .6 -7 Mach number, M

.8 -9 1.0

.40

.36

.32

28

2 4

.20

. / 6

. /2

.08

.04

0

-.04

-.08

-. 1 2

-. / 6 0 -1 2 .3 .4 3 .6 .7 .8 .9 1.0

Much number, M

(61 rr, = 0 4 Figure 19.- Conc/uded.

.08

4

0

0

0

.I .2 .3 .4 .5 .6 .7 A9 .9 LO Much number, M

Figure 20.- The voriution of increment of e/evcrfor hinge-moment coefficient due to tub defection with Much number. Q ~ . Oo; R, 2,000,000.

114

.. .

.. . ".

" . " -

. . . -. . . - . .

M=0.94

"-0.93

s M=O. 90 0;

b M"0.85 s 'r,

%

24 . 2 0

8 ./6 b

./2 L -08 2 .04

b o M=O.Z/

7 M=O. 80 . . -

M d . 6 0 . . - . . - " . . . . " -

4 4 108 1/2 : /6 -.20

-8 -4 0 4 8 /2 /6 2024 .. .

-24-2046 -12 -8 -4 0 4 Angle of dfuck,,, deg €/evafor def/ecfion, 88 , deg

v- Figure 21.- The cornpufed effecf of a sealed, infwnal, aerodynamic balance on fhe c

vuriafion of e/evufor hinge-momenf coeffckwf wiA4 mg/e of uftuck und wifh e/e- vufor def/ecfion. st, 0; R, 2,ooO,ooO.

. . . . . . . . . . . . . . . . .

I * .- -- . - - .

" c

0 '.I 2 .3 .4 .5 .6 .7 .8 9 LO

Mach number , M T

Figure 22.- The variations of pitching-moment parameter C and aetv@umic-cenfw % location with Mach number. 8, , 0'; R, P,ooO,ooO.

RAGA RM ~ 9 ~ 1 3

NACA RM ~ 9 ~ 1 3

./ .2 .3 .4 .5 .6 .7. B .9 1.0 Mach number, M

4

Figure 24.- The independent effects of leuding-edge rouMness und removal of the elevufor-nose sed on the vuriufions of /iff parum- efers CL, , CLse, and a&* with Much number. R, 2,000,000.

.

,040 8 E

,020 3 Minimum drag coefficient 3i k k

O 40

u Q

Q, 0 u *.

.4 C 0 t h

$35 8 B

4 L. E rough., sealed L 4 P)

$ 30

-2 c,

g 25 * 4

~ P o I I I I I ' l ' ' ' l I I I I ' ' I I I I 0 . I 2 .3 .4 3 6 .7 .8 .9 1.0

Mach number, M

-57 F F Fipre 25.- Thi? ind?pe&nf effecfs of hading-edge roughness and renwval of the elevator-

nose seal on ihe wiaiions of minimum drag coefficient and aerodynamic-center location with Mach number. 0"; at , 0'; R, ~,OOO,OOO. w P

E

. . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . .

tests model of aspect ratio 4 - unt digital library/67531/metadc58274/m2/1/high_res… ·...

Documents