The Effects of Blowing Over Various Trailing-edge Flaps on an NACA 0006 Airfoil Section, Comparisons with Various Types of Flaps on other Airfoil Sections, and an Analysis of Flow and Power Relationships for Blowing Systems

mass-flow rate of blowing air foot of span per section mass-flow coefficient, pocvo

P.SV + S(Pj - Po)

C section jet-momentum coefficient, J j ( p j assumed CI 90C equal to po except as noted) mass-flow rate of blowing air mass-flow coefficient, c Q

p.A.V * + A.(p

- po)

j jet-momentum coefficient, J J j J (pj assumed c , q O s , equal to po except as noted), see Appendix A coefficients in the equations for wind-tunnel wall corrections Cl.. .5 h height of test section, ft ’, 2 section lift, lift per unit span, lb/ft m section pitching moment, pitching moment per unit span, ft-lb/ft V

M Mach number, a

pressure, lb/sq ft P dynamic pressure, lb/sq ft q

P - Po

P pressure coefficient, 9 , r radius, in., or fraction of wing chord R Reynolds number based on the wing chord S height of the nozzle opening measured normal to the wing chord line at the minimum cross-sectional area of the nozzle, ft height of the nozzle opening at the exit of a convergent- se divergent nozzle, ft the reference wing area affected by the nozzle span, s q ft SW t airfoil thickness, ft =When used without subscript t, the symbols p, p, and T denote - - static pressure, static density and static - temperature, _ - .. . > = _ . _ _ respectively.

T absolute temperature,2 OR

v velocity, ft/sec

X chordwise distance, in. or ft

distance normal to the airfoil ch rd lin , in. or ft

Y coordinates for identifying the position of the nose of the Xf, Yf

trailing-edge flap, percent of wing chord (see fig. 7)

a section angle of attack, deg

flap effectiveness parameter, -

( a % ICz =za

ratio of specific heats, 1.4 for air

Y 6 angle of deflection of the trailing-edge flap, deg angle of deflection of the nose flap, deg 6n correction factor for atmospheric conditions different from A standard conditions,

( z s ’ l & )

mass density of air,2 slugs/cu ft P Sub scripts a ambient conditions i ideal angle of attack conditions in the jet at the exit of the nozzle max maximum 0 free-stream conditions std sea-level standard conditions t total conditions (i.e., isentropic stagnation conditions) U uncorrected 2See footnote 1, page 4 .

.

. . .

Superscripts

* conditions where M = 1.0

EXPERIMENTAL INVESTIGATION WITH A THIN AIRFOIL Tunnel, Model, and Apparatus Tunnel.- Because of the l i m i t a t i o n s of t h e auxiliary a i r supply f o r t h e Ames 7- by 10-foot wind tunnel, it w a s necessary t o modify t h e t e s t section of t h e tunnel t o accomodate a model with a reduced span.

Figure 1 shows t h e symmetrically spaced flow dividers which were i n s t a l l e d i n t h e tunnel t o provide a 4- by 10-foot test section.

Each d i v i d e r extended upstream about 13 f e e t and downstream 12 f e e t from t h e center l i n e of r o t a t i o n of t h e model. The 6-foot-diameter aluminum t u r n t a b l e s were supported f l u s h with t h e surfaces of the dividers, as shown i n f i g u r e 2, and were a l i n e d with, and connected t o t h e e x i s t i n g tunnel turn t a b l e s . Airfoil-shaped f a i r i n g s were used t o s h i e l d t h e model support s t r u c t u r e from t h e a i r flow between the flow dividers and the o r i g i n a l f l o o r and c e i l i n g of the tunnel t e s t section. These f a i r i n g s had t h e NACA 65,-415 a i r f o i l section and a 58.75-inch chord.

They were sup ported from t h e t u r n t a b l e s i n t h e f l o o r and c e i l i n g of t h e o r i g i n a l tun n e l and were arranged t o change angle of a t t a c k with t h e model. Pressure surveys i n t h e modified t e s t section indicated t h a t t h e flow between the dividers i n the 4- by 10-foot t e s t section w a s e s s e n t i a l l y uniform.

Calibrated s t a t i c o r i f i c e s on t h e w a l l s of t h e t e s t section approximately 6 f e e t upstream from t h e center l i n e of r o t a t i o n of t h e model were used t o i n d i c a t e free-stream s t a t i c pressure.

Model.- I n f i g u r e 2, t h e 4-foot-chord model i s shown i n s t a l l e d Yn t h e modified t e s t section. The basic a i r f o i l section of the model was t h e NACA 0006, modified t o accommodate t h e nozzle used with t h e a i r blow ing system and the various trailing-edge f l a p s . A d e t a i l e d view of the e x i t of the nozzle, which extended along the e n t i r e span of t h e model on t h e upper surface, i s shown i n figure 3. Some d e t a i l s of t h e plenum chamber and nozzle shape a r e shown i n f i g u r e 4 together with t h e 15-percent-chord nose f l a p . The s t e e l p l a t e s forming t h e nozzle could be positioned by means of 19 spacers and tightening screws located at 2-1/2-inch i n t e r v a l s along t h e span. The r a t i o of the cross-sectional area of t h e plenum chamber t o t h e nozzle e x i t a r e a w a s l a r g e enough t o ensure t h a t t h e velocity of flow i n t h e plenum chamber w a s negligible with respect t o t h e e x i t i n g velocity. (With a nozzle e x i t height of 0.053 inch, s/c = 0.00110, t h i s area r a t i o w a s about 20 t o 1.)

Details of t h e trailing-edge f l a p s a r e shown i n f i g u r e 5. Each of the f l a p s could be deflected and positioned independently of the wing.

A removable f a i r i n g which could be i n s e r t e d i n t h e nozzle e x i t w a s used i n conjunction with f l a p A t o form t h e t y p i c a l s i n g l e - s l o t t e d f l a p arrangement. (The coordinates for f l a p A a r e presented i n f i g . 4.)

The p l a i n f l a p s were designed t o d e f l e c t about t h e hinge points shown i n figure 5 . Each of t h e s e p l a i n f l a p s w a s designed s o t h a t it f a i r e d i n t o t h e unmodified a i r f o i l contour a t about t h e x/c = 0.75 s t a t i o n . Flap B provided t h e basic shape t o which various nose sections were f i t t e d t o form f l a p s C, D, and E. Flap B w a s symmetrical and was f0rme.d by s t r a i g h t l i n e s from t h e t r a i l i n g edge tangent t o t h e nose radius of t h e f l a p . A comparison of t h e profixes of f l a p s A, B, and C f o r the same f l a p deflection i s shown i n figure 6 t o emphasize t h e d i f f e r e n t f l a p contours presented t o the a i r e x i t i n g from t h e nozzle. The chord of f l a p A w a s 30 percent; f l a p s B and C w e r e 25-percent chord, and f l a p s D and E d i f f e r e d s l i g h t l y from 23 percent, depending on t h e location of t h e i r hinge points. Flap F provided a 13-percent-chord f l a p based on a t o t a l wing chord of 42.35 inches. This reduction i n wing chord w a s a r e s u l t of shortening t h e chord of t h e f l a p . Thus with f l a p F, t h e air f o i l section p r o f i l e deviated from t h e NACA 0006 p r o f i l e , t h e thickness based on t h e shortened chord w a s 6.8 percent, and t h e nose f l a p w a s 17 percent of the chord.

A f i l l e r block and an adjustable p l a t e were attached t o the main wing t o provide similar wing-flap junctures f o r a l l t h e p l a i n f l a p s ( f i g . 3 ) .

For a l l t e s t s with t h e p l a i n f l a p s deflected or undeflected, t h e gap between t h e end of t h e adjustable p l a t e and t h e f l a p w a s 0.1 percent of the wing chord, Chordwise pressure d i s t r i b u t i o n s were obtained from three rows 0 1 o r i f i c e s , one row a t t h e midspan, and a row 6 inches from each end of t h e span. Both s t a t i c - and total-pressure tubes Were i n s t a l l e d i n t h e plenum Temper chamber along t h e span t o measure pressures Of t h e i n t e r n a l flow- a t u r e s i n t h e plenum chamber were measured by shielded t h e ~ o c o u P l e s a t t h r e e spanwise s t a t i o n s .

Apparatus.- A variable-speed a i r compressor located outside of t h e wind tunnel w a s used as the source f o r the compressed air. The maximum pressure r a t i o s ( r a t i o of plenum-chamber pressure t o free-stream s t a t i c pressure) available with t h i s equipment were of the order of 1.7 t o 1.8.

A section of f l e x i b l e piping w a s included i n the ducting between t h e a i r compressor and the s t r u c t u r e supporting t h e model t o prevent any of t h e forces i n the ducting from a c t i n g on t h e scale system. An "0" r i n g s e a l w a s used i n t h e ducting approaching the model s o t h a t the angle of a t t a c k of the model could be varied without appreciable loss of a i r from t h e blowing system. The m a s s r a t e of air flow through t h e ducting was meas ured by a c a l i b r a t e d o r i f i c e meter i n s t a l l e d i n the l i n e between t h e s e a l and the compressor.

Test Methods Procedure.- Data were obtained for free-stream Reynolds numbers of

2.3, 3.3, and 4.0 million; the corresponding free-stream Mach numbers

were 0.082, 0.117, and 0.143. Air flow through the nozzle was varied from zero to the maximum values obtainable with the air compressor, and was expressed in terms of the mass-flow coefficient, CQ, and the jet- momentum coefficient, cCL. The rate of air flow measured with the orifice meter was used to calculate the mass-flow coefficient, CQ. In addition, measurements o'f the pressure and temperature in the plenum chamber were used to establish the reservoir conditions of the jet flow exiting from the nozzle to calculate the momentum coefficient, cp. Isentropic flow from the reservoir conditions in the plenum chamber to the nozzle exit and a static pressure in the jet at the exit equal to free-stream static pressure were assumed in order to calculate the momentum of the measured mass flow leaving the nozzle. Pressure measurements taken along the span in the plenum chamber were nearly equal for all except the lowest operat ing pressure ratios, and, consequently, it was assumed that the flow ejected from the nozzle was uniform along the span. Because of the limited pressure ratio available, and because of the range of nozzle heights tested, it was necessary to reduce the free-stream velocity from 160 feet

per second (R = 4 . 0 million) to 92 feet per second ( R = 2,3 million) for

some tests to cover the range of momentum coefficients of interest. The nozzle-height to wing-chord ratios quoted herein are "effective" values; that is, they were calculated from the isentropic flow relationships by the use of measured values of the pressure ratio, the flow coefficients, (CQ and cP) and the wind-tunnel dynamic pressure for a wide range of flow conditions. These values, in most cases, agreed very well with physical measurements of the nozzle height made with pressure in the nozzle. The effect of the maximum internal pressure forces on the nozzle was to increase the nozzle height by about 0.002 inch (s/c = 0.00004). This increase due to the internal pressure forces did not vary with changes in the nozzle-height to wing-chord ratio.

Lift measurements were made with the wind-tunnel balance system for each flap at the various free-stream Reynolds nwnbers. Data were obtained for each flap deflection with the nose of the flap in various positions relative to the nozzle exit (or, relative : o the fairing in the case of the single-slotted flap). These tests, or surveys, as they will be called herein, were made to establish the best position of a flap for purposes of further testing. The nozzle exit was sealed by the fairing for the tests with the single-slotted flap. The selected locations of the nose of the single-slotted flap are shown in figure 7(a) for each of the flap deflections tested. With the other flaps the surveys were made for vari ous blowing conditions. Extensive surveys were made with flap A, and the various selected locations for the nose of the flap are shown in A were arbitrar figure 7(b). Three categories of flap position for flap ily established for purposes of discussion: these are the extended, intermediate, and against-the-nozzle positions indicated in figure 7(b).

The reasons for testing the flap in these positions are discussed in a (Effect of flap position). Surveys were made with the following section plain flaps in order to determine the effect of vertical location of the flaps with respect to the jet. In these surveys.,the flap was moved longitudinally the small amount required to close the gap between the flap and the nozzle.

Two operating procedures for obtaining the data were employed: First the quantity of air exiting from the nozzle (i.e.> CQ or cp) was maintained constant and the angle of attack was varied. Secondly, the angle of attack was maintained constant while the nozzle flow was varied from high values of CQ or cp to zero. The hysterisis effect on the lift coefficient between increasing or decreasing nozzle flows was found to be negligible in the limited, but representative, number of tests conducted to evaluate this effect.

Corrections.- Corrections to the angle of attack, lift, and pitching moment were applied as follows using the method of reference 14:

a = uu + ClCZU + C2Cmu

cz = c3czu

cm = c4cmu + c c

2 , 0.400 0.353 With the modified tunnel, the ratio of the wing chord to test-section height was 0.400 for the model with each of the flaps except flap F. In the latter case, the ratio was 0.353. Blockage corrections for the condition with a blowing jet of air are unknown. However, on the basis of the blockage studies presented in reference 12 for a chord to height ratio of 0.32, it was assumed that the blockage was small for the chord to hei,ghtratios of the present tests. No further analysis of the change in the wind-tunnel wall corrections due to the effects of a blowing jet was made.

Test Result s The lift data are assembled according to an arbitrary grouping of the flaps, and include data with and without blowing.

The data with blowing over the flap are presented in two forms: (1) section lift coef ficient as a function of the angle of attack (for a given nose and trailing-edge flap deflection, and for various constant values of the section jet-momentum and the mass-flow coefficients),. and (2) the section lift coefficient as a function of the jet-momentum and the mass-flow coefficients (for a given nose and trailing-edge flap deflection and for various angles of attack). Representative moment and midspan pressure- distribution data are presented only for flap A. These typical pressure- distribution data should be of value for flap loading analyses as well as for their general aerodynamic interest. The test data from the investi gation are presented in figures 8 through 60. For convenience, an index to these data is presented in table I.

Single-slotted flap.- Data were obtained with the single-slotted flap for comparison with the data obtained with the blowing flaps.

Figure 8 presents the test data for various nose flap deflections (for a trailing-edge flap deflection of 50°), from which a nose flap deflection of 30° was selected as optimum for use in further tests of the single- slotted flap without blowing. The basic data for various trailing-edge flap deflections with this nose flap deflection, and also with the nose flap undeflected, are presented in figure 9.

Flap A.- Data showing effects of blowing with both the nose flap and the trailing-edge flap A undeflected are shown in figure 10. A limited amount of data with the nose flap undeflected is presented in figures 11 and 12. Figure 11 shows the effect of deflecting the trailing-edge flap 3 0 ' and 60° (in the extended position) without blowing and with a large amount of blowing. Figure 12 shows the effect of various amounts of blowing for one trailing-edge flap deflection ( 6 = 7 0 ' ) . The effects of deflecting the nose flap are shown in figure 13 for specified blowing quantities and trailing-edge flap deflections. These data were used to select a value for the nose flap deflection for use in the tests with blowing. A value of 3 5 ' was considered to be the optimum value and it was used, except as noted, in the tests with blowing. The effects of blowing on the lift coefficients for various trailing-edge flap deflec

tions are shown in figures 14 to 1-9 with the trailing-edge flap in

extended positions (and with the nose flap deflected 3 5 ' ) . Data obtained with the flap against the nozzle and for trailing-edge flap deflections of 50°, 60°, and TO0 are presented in figures 20 to 22.

The effects of sealing the wing-flap gap, when the flap was against the nozzle, are presented in figure 23.

An investigation of the effects of changes in the nozzle heights was made with flap A against the nozzle and the data are presented in figures 24 to 29.

In order to obtain some indication of the effect of blowing over various portions of the span of the flap, a brief investigation was made with various spanwise portions of the nozzle blocked off. The data are presented in figure 30.

Plain flaps B, C, D, E, F.- Except for a limited number of tests conducted with flap C with the nose flap undeflected, the tests with the plain flaps were conducted with the nose flap deflected 3 5 ' . The effect of deflecting flap B is presented in figure 31 and the effects of blow ing are given in figures 32 to 34. Similar data are presented for flaps C and D in figures 35 to 42. Data of this type were not presented for flap E because the flow over the flap at the larger flap deflections was separated even for the highest blowing quantities. The effect of deflecting flap F is presented in figure 43 and the effects of blowing are given in figures 44 to 46.

Pitching moments and pressure distributions with flap A.- Typical changes of the pitching-moment coefficient associated with changes of flap deflection, nozzle height, and blowing quantity are presented in figures 47 to 51. Representative wing-flap pressure distributions at the midspan of the model are given in figures 52 through 59 for flap A in both the extended position and against the nozzle.

Discussion of Test Results Definitions.- The test results to be discussed are summarized in figures 60 to 63. In the discussion herein of the various effects of blowing over the trailing-edge flap of a thin airfoil, three frequently used quantities are the critical momentum coefficient, the ideal angle of attack, and the increment of lift coefficient at the ideal angle of attack.

The critical momentum coefficient is defined as the value of the momentum coefficient at which a large change occurs in the slope (dc2/d~~),,~ and above which only small increases in are obtained with additional c2 increases in cp for a constant angle of attack and flap deflection.

The critical momentum coefficients presented herein were determined from the data for an angle of attack of Oo. Observations of the pressure distribution over the various flaps indicated, in general, that the flow over the flaps was attached at values of the momentum coefficient that were slightly lower than the critical momentum coefficient as defined herein.

Because of the combined effects of the nose flap, trailing-edge flap, and the blowipg quantity on the lift characteristics of a thin airfoil, difficulty was encountered in eference slope taken tor selecting an angle of attack airfoil without blowing suitable for comparing lift and with 8 = 0 : increments. In order to resolve this difficulty satisfactorily, the increment of lift coefficient (labeled (Acz)~ in sketch (a)) was measured at the largest neg ative angle of attack for which the lift curve was essentially linear. Pressure distributions indicated that at this angle no separation of the flow occurred

" ; ; g

on the lower surface of the air ideal" angle of attack foil with the trailing-edge flap deflected. This angle of attack is defined as the "ideal" angle Sketch (a) of attack, and the lift increments measured at this angle reveal the effects of changes in the blowing parameters and flap characteristics in a manner that is reasonably independent of interference from other factors.

One reason for this is that at the ideal angle of attack the pressure gradient on the upper surface of the forward portion of the airfoil is the most favorable that exists on the airfoil for any angle of attack for which there is no separation from the lower surface. The increment of lift coefficient was measured from the linearly extended lift curve for the model with the trailing-edge flap undeflected and with no blowing. It was necessary to extend this curve because the flow separation from the lower surface of the airfoil near the ideal ang,leof attack without blow ing produced a change in the slope of the lift curve which was otherwise constant for a wide range of angles of attack.

The experimental results are also compared with theoretical lift increments computed by the use of Glauert's relationship for a thin air foil with a hinged flap (ref. 15) y without consideration of the effects of blowing, but corrected for the effects of airfoil thickness ratio Effect of flap __.___ position.- Surveys were made to select the location of e a - a c h flap deflection. With the single-slotted flap, the locations of the flap were selected to provide the optimum lift character istics. Shown in figure 7(a) are the selected logations of the nose of

the flap for flap deflections of bo0, 50°, and 60 . It is apparent that

the optimum position of the nose of the flap was always below, and near the exit of the slot lip.

The selected locations for the nose of flap A are indicated in With the flap figure 7(b) for each of the specified flap deflections.

in the extended positions, the selected locations of the nose were determined from surveys conducted to determine the optimum lift character Thus, in figure 7(b), istics for a high value of the momentum coefficient.

the line connecting the points locating the nose of the flap represents the flap path required to obtain the optimum lift characteristics for a high value of the momentum coefficient. It is worthy of note that for flap deflections of 50' and above, and for the flap in either the extended or against-the-nozzle positions, the nose of the flap always protruded into the jet (see fig. 7 ( b ) ) . The surveys indicated that at these flap deflections the flow would not remain attached when the flap was removed from the jet. The effect of flap position is evident in the basic lift data (figs. 17 through 22) for the flap in the extended and against-the nozzle positions. Figure 60 (which includes the small amount of data for the flap in the intermediate positions) presents lift data for 0 ' angle of attack to provide a more direct comparison of the-effect of longitudinal position of the flap. It appears from figure 60 that the rate of change of critical momentum coefficient with increasing distance of the flap from the nozzle exit continually increased. For example, with the flap deflected 60°, moving the flap longitudinally 0.5-percent chord away from the nozzle doubled the critical momentum coefficient, and with the flap in the extended position, the critical momentum coefficient was increased approximately eight times. It can also be seen in figure 60 that the rate of change of the lift coefficient at the critical momentum coefficient with increasing distance of the flap from the nozzle exit was approximately constant.

The surveys with the plain flaps were made to determine the effect of vertical location of the flap with respect to the jet. The data 31 through 4 6 are for the optimum flap positions presented in figures which showed that the upper surface of the flap should be near the center of the jet. However, the effects of vertical position were found to be small so long as the upper surface of the nose of the flap was in the jet but below the upper surface of the airfoil contour. It should be noted that the hinge points for which the data are presented were shifted slightly from the design hinge points indicated in figure 5; the longi tudinal location was closer to the exit of the nozzle and the vertical location was shifted the small amount required to place the nose of the flap near the center line of the jet.

In considering the effects of flap position (and also the effects of flap profile presented in the following section), it should be remembered that in this investigation the velocity at the exit of the nozzle was subsonic and calculated with the assumption of isentropic expansion of the jet flow to free-stream static pr-ssure. With supersonic jet velocities, the question arises as to whether or not it would be desirable for a flap to protrude into the jet. However, consideration of the results of the present investigation which were obtained with subcritical pressure ratios, and those of reference 13 which were obtained with both sub- critical and supercritical pressure ratios, suggests that at least with plain flaps and convergent nozzles, the effects of flap position determined by the present investigation would be the same for pressure ratios up to moderate supercritical values.

Effect of flap profile.- The effects of flap profile are shown in figure 61 in which the iift coefficients at 0 ' angle of attack are given as a function of both the momentum coefficient and the mass-flow coeffi

cient. A study of the flap profiles (figs. 5 and 6) in conjunction with

these data indicates that the profile of the flap was of importance in securing a low critical momentum coefficient, but that the profile was of lesser importance for values of the momentum coefficient larger than the critical value.

For a given flap deflection (see fig. 6), the flaps whose profile enabled the exiting nozzle flow to be turned in a gradual manner had a lower critical momentum coefficient than the flap whose profile turned the exiting nozzle flow in an abrupt manner. Although both flaps A and C turned the air in a gradual manner, flap A had a lower critical momentum coefficient than flap C, particularly at the larger flap deflections. This may be due to the more gentle curvature of the profile of flap A compared to flap C (in the region away from the nose of the flaps), and it may also be due to the sharp nose shape of flap A, which projected into the jet close to the exit of the nozzle.

In addition to illustrating the effects of flap profile, the data

of figure 61 permit the effect of the ratio of flap chord to wing chord

to be estimated. This can be done by a comparison of the data for flap F (cf/c = 0.15) with the data for the other flaps (cf/c = 0.25 to 0 . 3 0 ) .

As a result of the design criteria for flap F (see the discussion in the section "Model") the profile of the flap was poor, resulting in a high critical momentum coefficient. From the previous discussion of the effects of flap profile it would appear that with a better flap shape, the high critical momentum coefficient could be reduced. However, the important point to note in figure 61 is that at high values of the momentum coefficient, where the effect of the profile has been shown to be of lesser importance, the lift-obtainedwith flap F compares favorably with that obtained with the flaps having larger ratios of flap chord to wing chord. This is evident particularly at the largest flap deflection, Thus, it may be true that, with blowing, the lift is relatively 6 = 7 0 ' .

insensitive to the flap-chord ratio.

Effect of changes in nozzle height.- The effect of changes in the r a t i o v t to wing chord on the lift increment at the ideal - angle of attack as a function of the momentum and the mass-flow coeffi cients was investigated using flap A in its position against the nozzle.

The results are presented for trailing-edge flap deflections of 50' and 6 0 ' in figure 62. The large reduction in the mass-flow coefficient, CQ, with reduction in the nozzle height for a given lift increment is In the range of nozzle height to wing-chord ratios from O.OOOl7 apparent.

to 0.00065, the effects of height-chord ratio on the lift increment for a given momentum coefficient were very small. In the investigation of reference 9 height-chord ratios in a low range (s/c = 0.00036 to 0.00072) were also tested, and the results showed no effect of changes in the nozzle height on the lift increment. Reference 13, which presents the results of a three-dimensional, full-scale investigation of the effects of the blowing air from a duct located in the flap of a swept-wing air also showed that the lift obtained at a given momentum coefficient plane, independent of the nozzle height for the range of values investigated was (ratios of nozzle height to mean aerodynamic chord between 0.00017 and 0.00067).

In the tests of the present investigation, however, an increase in the nozzle-height to wing-chord ratio from 0.00065 to 0.00110 resulted in a considerable loss in the lift increment obtained at momentum coefficients greater than the critical (see fig. 62), but there were no of nozzle height on the critical momentum coefficient significant effects at 0 ' angle of attack (figs. 20 through 2 9 ) . Data pertaining to the effects of nozzle height on the increment of lift coefficient obtained from reference 12 are shown in figure 62(~) for values of the height- These results show that increasing chord ratio from 0.0005 to 0.009.

from 0.0005 to 0.0015 brought about a much smaller loss in the lift s/c increment than that shown in the present investigation by changing s/c from 0.00065 to 0.00110. The marked effect of nozzle height shown 0.0015 to 0.0050 is question from in figure 62(~) for increasing s/c able because of changes that were made in the nozzle design and flap location. Since the limited amount of data presented herein indicates that the effects of changes in the nozzle height may depend partially on the particular nozzle and flap configuration used, the results obtained with flap A cannot be considered as general. However, for any particular blowing flap arrangement, the possibility of there being effects of nozzle height must be considered.

Effect of nose flap deflection.- Some of the effects of deflecting the nose flap are contained in the data of figures 12 and 1 3 for flap A, and in the data of figures 36 and 39 for flap C. The data obtained with the plain flap C were used to show the effects of nose flap deflection on the variation of the lift increment at the ideal angle of attack with momentum coefficient (fig. 63). The principal effect of deflecting the nose flap was to reduce the lift increment at small values of the momentum coefficient without affecting the critical momentum coefficient.

As the momentum coefficient was increased, the difference in the lift increment caused by deflecting the nose flap continually decreased, and at values of the momentum coefficient larger than about 0.16, a somewhat larger lift increment was measured with the nose flap deflected than with it undeflected. The greater lift increments with the nose flap deflected were due mostly to a difference in the lift-curve slopes of the base curves which were used in the measurement of the lift increments. This effect of the different lift-curve slopes of the base cmves was not significant at low values of the momentum coefficient because the ideal (The base curves were those obtained with angles of attack were small.

out blowing, with the trailing-edge flap undeflected, and with the nose flap either undeflected or deflected 35'. ) In the following sections, comparisons w i l l be made with the results of other investigations which employed airfoils having either no leading- edge device, or devices which differed from the nose flap of the present investigation. The data from the present investigation which will be used in the comparisons were obtained with the nose flap deflected.

Although this practice resulted in smaller lift increments in the low range of momentum coefficient, it is believed to provide a more realistic comparison because thin airfoils, such as the one of the present investi gation, would require some form of leading-edge device to delay leading- edge separation at high angles of attack.

-_ :. distribution pressure ~ Effect of blowing on the pitching moment and with flap A.- The data of figures 4 8 and 5l(a) typify, for the-flap in the extended and against-the-nozzle positions, respectively, the large changes that occur in the pitching moment as the momentum coefficient increases. However, as shown in the following table, the change in the pitching-moment coefficient due to a unit change in the lift coefficient was not significantly affected by blowing over the flap for either posi tion of the flap. The values of the momentum coefficients are larger than the critical momentum coefficient in each instance.

- __ _- - - - -_ Flap position Extended -1Againf-L the nozzle 6 35O 50° 6 0 ' 50° 60' ~

-

Cp 0 0.12 - 0 _ _ 0.03 o 0.03 . -.201-.22 -.26!-.22 -.22- -.22 . ~ 1 9 . - . 2 0 -.18 -.19 A T

" y

~ The very great differences that occur in the pressure distributions for the no-blowing and for the high-quantity blowing cases are clearly 52 to 59. When the jet attached to the flap, shown by the data of figures of the flap and the pressure a low pressure peak developed over the nose coefficient near the trailing edge became positive in value (e.g., see figs. 55 and 58). Note that a positive pressure coefficient on the nose of the flap exceeding a value of 1 . 0 is indicated in figures 52(b) and (c) for the 75.10-percent-chord station. These high positive pressures on the nose of the flap result from the direct impingement of the jet on the flap and occurred with the flap undeflected or deflected in its position against the nozzle.

COI"ARIS0NS AND EVALUATION OF TRE EFFECTS OF BLOWING ON LIFT The following comparisons of the effects of blowing on lift for the blowing-flap arrangements of the present and the referenced investiga tions are made in terms of quantities believed to be of most significance for the evaluation of relative flap effectiveness. These quantities are (1) the increment of lift coefficient at the ideal angle of attack, (2) the critical momentum coefficient and the increment of lift coeffi cient which was obtained at the critical momentum coefficient, (3) the rate of change of increment of lift coefficient with momentum coefficient (ac2ildc,)ai ,6 for values of the momentum coefficient which were

greater than the critical value, and (4) the momentum coefficient required

to obtain a lift increment equal to the theoretical increment of lift coefficient due to flap deflection without blowing. These quantities should be considered together, not individually, in order to form a complete picture of the relative lift effectiveness of blowing-flap arrangements. The airfoils of the referenced investigations were thicker than the airfoil of the present investigation and included types with and without leading-edge devices. It should be noted that differences exist in the value of the ratio of flap chord to wing chord for the various flaps of the present investigation as well as for the flaps of the refer enced investigations (see fig. 64). Unfortunately, sufficient data are not contained in the reports of these investigations 'to clearly establish the effects of changes in the ratio of flap chord to wing chord.

Lift-Coefficient Increment at the Ideal Angle of Attack Ln comparisons of the lift effectiveness of high-lift devices, the increment of lift coefficient obtained at a given angle of attack is usually presented as a function of the deflection of the device. This convention has been retained for the comparisons presented herein of the various arrangements of the flap and blowing system. However, an addi tional quantity, the jet-momentum coefficient has been included to show the effects of various amounts of blowing. The data of the present

investigation and of references 4, 5 , 9, and 12 (see fig. 64 for

sketches showing the various arrangements of flaps and blowing-system nozzles) are summarized in this form in figures 65 through 71. The increments of lift coefficient presented herein for the present investi gation were measured at the ideal angle of attack. The increments presented for the referenced investigations were measured at 0 ' angle of attack instead of at the ideal angle of attack because of insufficient data to define the latter angle. However, because the increment at 0 ' angle of attack was the largest that could be measured, and because Lt was thought that it would be essentially the same as that increment which would oceur at the ideal angle of attack, it was decided for the I I purposes of this report to refer to the increment of lift coefficient for

the referenced data as (AC~)~. Included in figures 65 through 71 are

theoretical increments of lift coefficient due to flap deflection without blowing and, also, increments which have been obtained with conventional high-lift devices such as single and double slotted flaps. Because of the small amount of published data for these devices on airfoils having the same thickness ratios and the same ratios of flap chord to wing chord as the airfoils considered herein, it is difficult to make comparisons of these devices with all of the blowing-flap arrangements; thus, only

data from the present investigation and from references 16 and 17 are

considered. Consequently, these data for the single and double slotted flaps were included in these figures only where it was thought that comparisons with the blowing data would have some validity and interest.

The lift-coefficient increments obtained at the ideal angle of attack with the various blowing-flap arrangements on the thin airfoil of the

present investigation are shown in figures 65 through 67; those obtained

for the airfoils of the investigations of references 5 , 9, 4, and 12, for which the airfoil thickness-chord ratios were 9, 10, 12, and 15 percent, respectively, are shown in figures 65 through 71.

It is evident from even a cursory examination of figures 65 through 7lthat large differences exist among the various airfoils and blowing- to a given amount of blow flap arrangements in regard to their response ing, and that with a sufficient amount of blowing the theoretical incre ments of lift coefficient were exceeded. A study of these figures reveals that with a given momentum coefficient an increment of lift coefficient could be obtained with the 6-percent-thick airfoil that equaled, or exceeded, the values obtained with the thicker airfoils of the referenced investigations. The data indicate that for some of the configurations additional lift effectiveness could be expected for flap

deflections above 6 0 ' or 7 0 ' . This is particularly evident from the data

small nozzle for the thin airfoil of the present investigation with the heights (see figs. 66(a) through 66(d)).

Critical Momentum Coefficient and Increment of Lift Coefficient

Presented in figure 72 is the variation of the critical momentum

coefficient with trailing-edge flap deflection for the data from the present investigation and from the referenced investigations. As shown in this figure, the critical momentum coefficient generally increased with increasing flap deflection and with movement of the flap away from the nozzle exit. This increase with flap deflection was small in some cases but very rapid in others. The increase with movement of the flap away from the nozzle exit is shown by comparing the results for flap A in its position against the nozzle and in the extended position. The critical momentum coefficients obtained with flap A in its position against the nozzle were smaller than those measured for any of the blowing-flap arrangements of the referenced investigations and did not exceed a value of about 0.03 for flap deflections up to TO0.

The increments of lift coefficient obtained at the critical momentum coefficients corresponding to those given in figure 72 are presented in

figure 73 together with the theoretical lift increments due to flap

deflection without blowing. A n inspection of these two figures shows that there were large variations in the critical momentum coefficient and in the lift-coefficient increments measured at the critical momentum coefficient for the various blowing-flap arrangements. The differences between the measured lift increments and their corresponding theoretical lift increments also varied widely. For example, at 60° flap deflection the largest critical momentum coefficient for the data of the present investigation was about eight times greater than the smallest value, and the increments of lift coefficient varied from about 60 to 99 percent of their theoretical values. At first thought it might be expected that such differences in the increments of lift coefficient should not occur because, for the critical momentum coefficient, separation of the flow over the flap was prevented. Control of separation of the flow over the flap, however, is a necessary but not a sufficient condition for attain ment of the theoretical lift increment. In addition, the amount of blow experimental case must be controlled t9 provide a circulation ing in the strength around the airfoil equivalent to that of the potential flow solution. Since the amount of blowing required to prevent separation of the flow differed greatly for the various flaps, the circulation strengths, and hence the resulting lift increments, also differ greatly.

It is apparent from the preceding discussion and example that in evaluations of the relative lift effectiveness of blowing-flap arrange consideration must be given to both the critical momentum coeffi ments, cient and to the increment of lift coefficient obtained for the critical momentum coefficient,

Examination of figures 72 and 73 shows, from the results of the

present investigation, that the critical momentum coefficient and the associated increment of lift coefficient were unchanged for nozzle-height to wing-chord ratios of 0.00065 or less. They were also unchanged for the height-chord ratios of 0.00036 and 0.00072 which were investigated in reference 9 . The data from reference 12 show a large effect of height- chord ratio, and the results obtained with the smallest nozzle heights indicated characteristics that differed from those obtained with the larger ones. It appears, therefore, that the effects of changes in the nozzle-height to wing-chord ratio are small for small values of this ratio (say, for values of less than 0.001), but may be significant s/c for larger values (say, for s/c greater than 0.001).

Rate of Change of Increment of Lift Coefficient With Momentum Coefficient The rate of change of the increment of lift coefficient with momentum coefficient (Uc,i/dcp)q, &, measured at values of the momentum

74 as a

coefficient greater than the critical, is presented in figure function of flap deflection-forthe flaps of the present and the refer A large value of ( ~ c ~ ~ / d c ~ ) , ~ , ~ is, of course, ence investigations.

desirable, but 'the significance of this parameter in assessing relative flap effectiveness depends also upon the critical momentum coefficient and the increment of lift coefficient at the critical momentum coefficient.

The effects of changes in the nozzle-height to wing-chord ratio on (dn~2~/dc~),~,~ were very small for flap A of the present investigation, but were large for the flap arrangement of reference 12, which had a much larger variation in the nozzle height. A considerably higher slope was measured for flap A in its position against the nozzle compared to that obtained in its extendbedposition. It is of particular interest to note the superiority of plain flap C, which was hinged on the lower surface, compared to plain flap B, which was hinged on the airfoil center line.

There was no marked effect of airfoil thickness ratio on (dnc,i/dcp)q,G as evidenced by the fact that this parameter was as large, in general, for the various flaps on the thin airfoil of the present investigation as it was for the flaps on the thicker airfoils of the referenced investigations.

Momentum Coefficient for Theoretical Increment of Lift Coefficient The value of the momentum coefficient required to achieve the theoretical lift increment is presented in figure 75.3 The accuracy of measuring the momentum coefficient required to achieve the theoretical lift increment depends to a great extent upon the rate of change of the lift increment with momentum coefficient (dnc, Although the absolute value of the momentum coefficient in a particular case may be

difficult to determine accurately, the values shown in figure 75 were all

obtained in a similar manner providing a common basis for comparison.

In general, the values of the momentum coefficient required to attain the theoretical increment of lift coefficient with the 6-percent thick airfoil were of the; same order of magnitude as those measured f o r 3A similar presentation has been noted in reference 18. The larger values of the momentum coefficients presented herein are due to the inclusion of the airfoil thickness correction in computing the theoreti cal lift increments as previously mentioned.

- thicker airfoil sections. In view of the variety of the blowing-flap arrangements considered, the data show very similar trends as a function of flap deflection, with but one exception - the data of reference 5 .

For this flap it is believed that the long overhang of the upper surface of the nozzle (see fig. 64) and the large distance from the nozzle exit to the flap resulted in a particularly poor blowing-flap arrangement. The advantages of the small nozzle-height to wing-chord ratios are evident from the reference data as well as the data of the present report. The values of the momentum coefficient required for the theoretical lift increment for values of less than 0.00065 were not determined in s/c the tests of the present investigation because of limitations of the available pressure ratio. However, on the basis of an examination of the limited amount of data available, no significant changes in the required momentum coefficient would be expected for the range of values of S/C from 0.00065 to 0.00017.

The data of figure 75 indicate that flap A in the extended position

required a smaller momentum coefficient to achieve the theoretical lift increment than it did in its position against the nozzle. In practical applications where the available momentum coefficient may be limited, the small value of the momentum coefficient required to achieve the theoreti cal lift increment probably would not be as important as the undesirable large value of the critical momentum coefficient that occurs with the flap in the extended position. Flap F had a flap-chord to wing-chord ratio of 0.15 compared with 0.25 to 0.30 for the other flaps considered.

Thus, the theoretical lift increment for flap F was smaller than for the other flaps. As previously shown (see fig. 61) the lift coefficients obtained (for momentum coefficients greater than the critical) with flap F This combination compared very favorably with those of the other flaps.

of a smaller theoretical lift increment and the relatively good flap effectiveness resulted in a considerably smaller momentum coefficient required to achieve the theoretical lift increment for flap F compared to those of the other flaps of the present investigation. The superiority of plain flap C in this regard compared to plain flap B was due to a larger value of (dAc2i/dcII)ai,6 obtained with flap C, since the critical momentum coefficients and the lift increments at the critical momentum coefficient were practically the same for these two flaps.

THEORETICAL FLOW AND POWER RELATIONSHIPS Flow Relationships The basic flow coefficients of interest for a blowing system are the mass-flow coefficient, CQ, and the jet-momentum coefficient, cP.

Figures 7 6 and 77 are presented to show the theoretical relationship

among these coefficients and the operating pressure ratio, the ratio of 2 1 nozzle height t o wing chord (proportional to Aj/S, for the three- dimensional case), and the free-stream Mach number. Appendix A presents the derivation of the equations upon which the figures are based. The

chart of figure 7 6 is applicable only where the pressure ratio is less

than the critical. The chart of figure 77 present's the relationships

for pressure ratios as high as 10, based on isentropic flow with an ideal nozzle.

It is to be noted that the definition of the jet-momentum coefficient is based on the assumption that the mass flow leaves the nozzle exit with the velocity that would be obtained by full isentropic expansion to free- stream static pressure. However, it should be realized that the momentum coefficients calculated on this basis do not always represent the true total momentum of the flow at the exit. A difference between the actual and the computed value of the momentum coefficient occurs when the exit or when the pressure is not equal to the free-stream static pressure, pressure ratio is supercritical and differs from the "design" value. The magnitude of the difference which may occur for pressure ratios above the critical is evident from the ratio of the jet-momentum coefficient for a convergent nozzle to that for a convergent-divergent nozzle for isentropic flow. The variation of the ratio of these momentum coefficients with pressure ratio is shown in figure 7 8 for pressure ratios less than 10.

The derivation of the relationship is presented in Appendix A. It is apparent that as the pressure ratio increases, the ratio of the momentum coefficients decreases until, at a pressure ratio of 10, the jet-momentum coefficient that could be obtained with a convergent nozzle is 0.93 of that which could be obtained with a convergent-divergent nozzle.

A unique solution of the two equations shown in figures 76 and 77 is

obtained by drawing a rectangle, such as the ones shown in these figures.

The rectangle connects equal values of free-stream Mach number in the upper and lower halves of the figure with the corresponding values of cp and s/c for the associated values of CQ and pressure ratio. For a particular solution, two of the parameters, in addition to the Mach number, must be ~pecified.~ A sequence of changes must occur among the various parameters shown in the figures whenever a change occurs in the value of any one of them. In the following examples the use of the charts is demonstrated. In general, certain changes dependent on the free-stream Mach number must occur in the values of the various parameters if the free-stream Mach number is changed. For example, consider the chart of figure 7 6 which applies for the range of subcritical pressure ratios.

I f the momentum coefficient and the nozzle height remain constant and the free-stream Mach number is changed, the mass-flow coefficient remains _ _ 4The lines of constant dynamic pressure', qo

(figs. 7 6 and 77), are

based on an absolute free-stream total pressure equal to Pstd, and they would be changed for other free-stream conditions. These lines are included in these figures for their general usefulness in problems con cerned with sea-level atmospheric wind tunnels.

constant and the pressure ratio must change. Thus, assume the initial conditions indicated by the dashed rectangle (i.e., cp = 0.06;

s/c = 0.0007; M, = 0.10; Pt /po = 1.325; and CQ = 0.0047). Now assume the

j

free-stream Mach number is increased to 0.14. By the process of succes

sive approximations the required rectangle closure yields the results that the pressure ratio would have to increase to 1.73, and CQ would remain the same. The fact that the mass-flow coefficient is invariant with free-stream Mach number for subcritical pr'essure ratios and for the conditions typified by this example (i.e., for a constant cp and s/c)

can be proved by differentiating the equations shown in figure 76. For

supercritical pressure ratios the mechanics of solving the equations shown

in figure 77 are identical to those indicated above for the subcritical

that is, the required closed rectangle must be determined.

pressure ratios; With the assumption of the initial conditions indicated by the dashed rectangle in figure 77 (cp = 0.08; s/c = 0.00057; Mo = 0.14; Pt./po = 2.35; J and CQ = 0.0048), a change in free-stream Mach number to 0 . 2 0 increases the pressure ratio to 3.85 and CQ increases to 0.0053. For the range of supercritical pressure ratios the derivatives of the equations shown

in figure 77 indicate that with a given momentum coefficient and nozzle

geometry, the mass-flow coefficient will vary with free-stream Mach number.

The preceding examples indicate how blowing-system data for particular free-strearh Mach numbers can be properly modified and adapted for use at other free-stream Mach numbers.

The inserts in figures 7 6 and 77 showing typical scale changes are

included to indicate the manner in which the range of values of cp, CQ, and s/c can be modified, provided the range of values of free-stream Mach number and the pressure ratio remain the same. With this provision the values of cp, CQ, and s/c can be multiplied or divided by powers of 10 as desired.

Power Relationships The power required to operate a blowing system can be used as a basis for comparing various arrangements of a flap and blowing system.

In Appendix B a power relationship is developed which is convenient for use in such comparisons. The final equation (eq. ( B 5 ) ) relates the section mass-flow coefficient, frge-stream Mach number, and pressure ratio, to the horsepower required per square foot of wing reference area. This horsepower relationship is based on the assumption of isentropic compres sion from free-stream total pressure to the jet total pressure, and is

shown in figures 79 and 80 for pressure ratios up to 1.9 and 10, respec

tively. It should be noted that the pressure ratio in these figures ptj/pto differs from the pressure ratio, pt /po which is given in the j flow charts. The lines of constant dynamic pressures shown in these figures are subject to the restrictions noted in footnote 4.

As an illustration of the application of the power and the flow charts, a comparison of the horsepower per square foot of wing reference area, the mass-flow coefficients, and the pressure ratios theoretically required at the value of the critical momentum coefficient for several of the arrangements of the flap and blowing system previously discussed is presented in figure 81. The value of the critical momentum coefficient for each arrangement and the corresponding lift increments have been

presented in figures 72 and. 73, respectively. I t is evident from

figure 81(a) that at a given Mach number there was a large variation in the power requirements for the various arrangements, and in some cases there were large effects of flap deflection. In general, there was an increase in the power required with an increase in Mach number, and the magnitude of the increase varied greatly among the various arrangements.

I f the air is provided by auxiliary compressing equipment, the power required is of greatest importance in the design of a blowing system.

However, if the air is supplied by bleeding from a jet engine, the mass flow, or cQ, is the more important quantity (fig. 81(b)). A large vari ation in the values of the mass-flow coefficients for the various flaps and blowing systems was evident, although for any particular case CQ was invariant with Mach number. Figure 81(c) shows that the required pressure ratio generally increased with increasing Mach number, and, also, that at a given Mach number there was a large variation among the various arrangements. The advantage, from the standpoints of power and mass-flow coefficient, of positioning the flap against the nozzle and using small nozzle heights is apparent throughout the comparisons afforded by figure 81.

CONCLUDING REMARKS The present report consists of (1) an experimental investigation made to determine the effects of blowing a jet of comparatively low- pressure air from a duct in the main portion of the wing over various types of trailing-edge flaps on an NACA 0006 airfoil, (2) a comparison and evaluation of the effects of blowing on lift, using the results of investigation and those of previous Investigations, and the present (3) an analysis of the theoretical flow and power relationships of a blow ing system.

, Tests of flap A in various positions with respect to the nozzle showed that (1) the nose of the flap should protrude into the exiting nozzle flow, and (2) the critical momentum coefficient, and the lift obtained at the critical momentum coefficient, decreased as the gap between the flap and the wing was reduced.

Tests of flaps having different profiles indicated that the flaps whose profile enabled the exiting nozzle flow to be turned in a gradual manner had a smaller critical momentum coefficient than the flaps whose profile turned the exiting nozzle flow in an abrupt manner.

. . .. .-_...

The lift obtained with blowing over a 15-percent-chord flap compared favorably with 23- and 30-percent-chord flaps at the higher values of the momentum coefficient. The critical momentum coefficient was large with the short chord flap but it could probably be reduced by changes in the flap profile.

Tests on flap A indicated that the effects of nozzle height on the increment of lift coefficient obtained for a given momentum coefficient were small in the range of nozzle-height to wing-chord ratios from O.OOOl7 to 0.00065. A further increase in the nozzle-height to wing- chord ratio to 0.00110, however, showed a considerable loss in the lift increment. There were no significant changes in the critical momentum coefficient with changes in the nozzle height.

The change in the pitching-moment coefficient due to a unit change in lift coefficient was not significantly affected by blowing.

Comparison of the data for the thin airfoil of the present investiga tion with other data for thicker airfoils and somewhat different blowing- (1) the increments of lift coefficient flap arrangements showed that obtained for a given momentum coefficient with the thin airfoil were comparable with, or exceeded, those values obtained with the thicker air foil sections; (2) flap A positioned against the nozzle had smaller critical momentum coefficients than the flap arrangements used with the thicker airfoils; (3) the rate of change of the increment of lift 'coef (measured above the critical value) for ficient with momentum coefficient the thin airfoil was comparable to that of the thicker airfoils; and

(4) the momentum coefficient required to attain the theoretical increment

of lift coefficient with the thin airfoil were of the same order of magni tude as those measured for the thicker airfoil sections.

A theoretical study was presented which established the relationship among the air flow and power parameters applicable to the general blowing case. Charts were presented showing these relationships. With the aid of these charts an analysis was made to show the magnitudes of the flow and power parameters for several blowing-flap arrangements operating at their critical momentum coefficients, and also, to show the effect of changes in the free-stream Mach number on these parameters. It was found that the horsepower per square foot of wing reference area, and the pres sure ratio, increased with increasing Mach number, but that the mass-flow coefficient remained constant when the pressure ratio was subcritical.

Ames Aeronautical Laboratory National Advisory Committee for Aeronautics Moffett Field, Calif., Mar. 1, 1956 (Reissued by Ames Research Center, National Aeronautics and Space Admin istration, Moffett Field, Calif., Jan. 13, 1976.)

APPENDIX A DERIVATION OF THE EQUATIONS RELATING THE GEOMETRIC AND AIR-FLOW PARAMETERS FOR A BLOWING SYSTEM In the subsequent development of the various relationships involving the mass-flow coefficient, the jet-momentum coefficient, and the ratio of nozzle area to wing reference area (proportional to s/c for the two- dimensional case), it is assumed that the nozzle flow is for a perfect gas, that the flow is uniform, and that the cOmpression from free-stream total pressure to the jet total pressure is isentropic.

By definition, the jet-mass-flow coefficient is For adiabatic flow conditions and for y = 1.4, this equation becomes For the assumption of isentropic compression between the free stream and the jet reservoirs, (A3 and, in general, a 7-1

pt = p(l + 0.m )

then the mass-flow coefficient becomes In application, equation (A5) must be modified to suit particular condi tions. With an ideal nozzle, complete expansion of the flow occurs to pressure po so that pj = po. Also, for pressure ratios greater than critical, the ideal nozzle must be convergent-divergent and for pressure ratios less than critical the nozzle must be convergent. Thus, for an ideal nozzle, and ptj/po greater than critical, (note that Aj/A* and Mj are functions of pt./po and their values are J readily obtainable from tables such as those in reference 19). For the two-dimensional case, the section mass-flow coefficient becomes Also, for the ideal nozzle, and pt./po less than critical, J Aj Mj

C Q = - - s , Mo

or, for the two-dimensional case the section mass-flow coefficient is With a convergent nozzle and pressure ratios greater than critical, the static pressure in the jet at the exit of the nozzle will not equal the

free-stream static pressure (p. # po), and the Mach number of the jet at

J the exit of the nozzle will be 1.0.

By use of equation (Ab) in (A5), the jet-mass-flow coefficient becomes where Mj = 1.0. As would be expected, equations (A6) and (A8a) provide equal values of CQ at equal values of pt./po, if Aj/& for the con- J vergent nozzle equals A * / & for the convergent-divergent nozzle. For the two-dimensional case the section mass-flow coefficient is By definition, the jet-momentum coefficient is total momentum of the --__ .. flow at nozzle exit qosw with the relationship

90 = $ P$O"

becomes equation (~9)

2 Aj [" P' (1 + yMj2)

---

-

cv yMo2 S, Po

- - po, then for both subcritical and If the nozzle expansion is to pj supercritical pressure ratios Mj2 Aj C y = 2 - - Mo2 S, Combined with equation (A5), equation (A12) becomes for the case of isentropic flow For the two-dimensional case the section jet-momentum coefficient is By the use of equation (All) a comparison can be made of the total momentum at the exit of an ideal convergent-divergent nozzle with that at the throat (which would be the total momentum for a convergent nozzle). Thus F !

I n t h e i s e n t r o p i c case f o r P j - - Po, and using equation ( A b ) , or

" + " 1 .?h8fn. */-n-\ - 1

(Note t h a t (Pt */p0) = (Ptj/Po), and t h a t both A*/Aj and M j a r e a func t i o n of ( p Thus, equation ( A 1 6 ) gives t h e r a t i o of t h e t o t a l momentum a t t h e e x i t of a convergent nozzle t o t h a t a t t h e e x i t of an i d e a l convergent-divergent nozzle having t h e same t h r o a t area as t h e convergent nozzle.

The c h a r t s of figures 76 and 77 present a graphic solution of t h e equations i n t e r r e l a t i n g t h e mass-flow c o e f f i c i e n t , free-stream Mach number, t h e momentum c o e f f i c i e n t , t h e r a t i o of nozzle a r e a t o wing reference a r e a (proportional t o s/c f o r t h e two-dimensional c a s e ) , and t h e pressure r a t i o . For a nonisentropic process between t h e r e s e r v o i r s of t h e f r e e stream and t h e j e t , it i s necessary t o t a k e i n t o account t h e changed r e s e r v o i r conditions of t h e nozzle flow. It should be noted i n connection with t h e s e c h a r t s t h a t t h e t h e o r e t i c a l momentum of t h e j e t may d i f f e r consid erably from t h e a c t u a l value. For example, t h i s occurs when t h e pressure f i e l d i n t o which t h e j e t exhausts from t h e nozzle i s l e s s than t h e f r e e - stream s t a t i c pressure. Then t h e nozzle flow i s subject t o an e f f e c t s i m i l a r t o t h e Coanda e f f e c t f o r a j e t exhausting i n t o ambient air; t h a t i s , t h e a c t u a l pressure at t h e e x i t of t h e nozzle i s reduced below t h e free-stream s t a t i c value, thereby increasing t h e e f f e c t i v e pressure r a t i o .

f o r pressure r a t i o s less than c r i t i c a l , a reduced nozzle-exit pres Thus, sure would increase t h e m a s s flow and t h e momentum of t h e j e t above t h e computed f o r a pressure r a t i o based on t h e free-stream values t h a t would be s t a t i c pressure. For pressure r a t i o s above t h e c r i t i c a l t h e r e would be no e f f e c t on t h e m a s s flow, b u t t h e momentum of t h e j e t would increase with an increase i n t h e e x i t v e l o c i t y . For pressure r a t i o s l e s s than c r i t i c a l t h e l o c a l pressure f i e l d at t h e e x i t of t h e nozzle i s u s u a l l y unknown, o r d i f f i c u l t t o obtain, so t h a t it i s much more convenient t o base t h e momen t u m c o e f f i c i e n t on t h e free-stream s t a t i c condition; t h i s w a s t h e case i n t h e present r e p o r t . For pressure r a t i o s above t h e c r i t i c a l t h e l o c a l pressure f i e l d should only have a s m a l l e f f e c t on t h e o v e r - a l l pressure However, as equation (A16) i n d i c a t e s , t h e momentum of t h e j e t w i l l r a t i o .

depend on t h e nozzle design. Thus, p a r t i c u l a r l y at pressure r a t i o s much g r e a t e r than c r i t i c a l , t h e computation of t h e momentum c o e f f i c i e n t should be i n accordance with whether t h e nozzle i s convergent, or convergent- divergent.

APPENDIX B DERIVATION O F THE POWER REQUIRED TO COMPRESS THE AIR FOR A BLOWING SYSTEM I n a steady-flow process t h e power required t o maintain t h e flow i s defined as t h e product of t h e m a s s flow and t h e work done per u n i t of mass flow. For isentr‘opic flow r e l a t i o n s h i p s t h e horsepower required t o compress t h e blowing-system air from free-stream t o t a l pressure t o t h e j e t t o t a l pressure i s S u b s t i t u t i n g equation ( A l ) i n t o (Bl) and expressing t h e v e l o c i t i e s and d e n s i t i e s i n terms of Mach number, t o t a l pressure, t o t a l temperature, and stagnation v e l o c i t i e s of sound y i e l d s t h e following equation f o r t h e horsepower per square foot of wing reference area expressed i n terms of t h e section mass-flow c o e f f i c i e n t (B2) 1/ 2 equation With equation ( A 3 ) , and noting t h a t (ato/astd) = (Tto/Tstd) ( B 2 ) becomes L Regrouping t h e terms t o provide t h e pressure r a t i o ptj/pto within t h e bracketed expression gives

034)

Equation (B4) i s applicable for use i n f l i g h t or atmospheric wind tunnels.

However, t h e total-temperature r a t i o and t h e t o t a l - p r e s s u r e r a t i o must be

\

evaluated d i f f e r e n t l y i n each application. If h i s a correction f a c t o r f o r ambient or atmospheric conditions d i f f e r i n g from standard,

A = [(&y2(&J]

and by t h e use of t h e approximation t h a t (1 + 0.2M02) = 1.0 i n equa

t i o n ( B k ) , t h e corrected horsepower per square foot of wing area becomes A graphical solution of t h i s equation i s presented as f i g u r e s 79 and 80.

With t h e assumption t h a t t h e Mach number function equals 1.0 t h e r e r e s u l t s i n t h e horsepower per square f o o t of wing area of about a maxi" e r r o r 1 and 3 percent f o r pressure r a t i o s up t o 10 f o r t h e f l i g h t , and f o r t h e wind-tunnel solutions, respectively. It w i l l be noticed t h a t t h e t o t a l 7- 1

pressure r a t i o i n equation (B5) (pt./pt o)T could be put i n t h e form

Y - 1 '\ J

@tj/p+T[l/(l + O.2Mo2], but i n this case t h e assumption t h a t

(1 + 0.2M02) = 1.0 r e s u l t s i n increasingly l a r g e e r r o r s a s t h e pressure

r a t i o approaches 1.0. 79 or 80 t o f i n d t h e horse Thus, i n using f i g u r e s The power function, t h e t o t a l - p r e s s u r e r a t i o P t j / p t o must be used.

flow c h a r t s of f i g u r e s 76 and 77 give t h e pressure r a t i o i n terms of f o r t h e given Mach number pt./po, which must be multiplied by po/pto J f o r use with t h e horsepower charts. The constan-t! "q" t o f i n d ptj/pto l i n e s on t h e s e power c h a r t s a r e r e s t r i c t e d t o wind-tunnel usage f o r t h e same reasons discussed i n footnote 4 i n regard t o t h e flow c h a r t s .

1, Seewald, F.: Increasing Lift by Releasing Compressed Air on Suction Side of Airfoil. NACA T M 441, 1927.

2. Reid, E . G., and Bamber, M. J.: Preliminary Investigation on Bound ary Layer Control by Means of Suction and Pressure with the U.S.A.

27 Airfoil. NACA T N 286, 1928.

Wieland, K.: Experiments With a Wing From Which the Boundary Layer is Removed by Pressure or Suction. NACA TM 472, 1928.

Schwier, W.: Lift Increase by Blowing Out Air, Tests on Airfoil of 4 .

12-Percent Thichess , Using Various Types of Flap. NACA TM 1148, Schwier, W.: Lift Increase Produced by Blowing a Wing of a Profile 5 .

Thickness of 9 Percent, Equipped With a Slat and a Slotted Flap.

Rep. No. F-TS-645-RE, Air Materiel Command Trans., Aug. 1946.

6. Boyer, Luther J.: Preliminary Investigation and Evaluation of the

Coanda Effect. Tech Rep. No. F-TR-2207-NDYAir Materiel Command, 1948.

Aug.

Nunemaker, John J., and Fisher, Jack W.: Two-Dimensional Wind Tunnel hvestigation of Boundary-Layer Control by Blowing on an NACA 23015 Airfoil. Rep. No. 023, Municipal Univ. of Wichita kgr., Apr. 1950.

8. Rebuffet, P., and Poisson-Quinton, Ph.: Investigations of the Boundary-Layer Control on a F u l l ScaPe Swept Wing With Air Bled Off from the Turbojet. NACA TM 1331, 1952.

Harkleroad, E. L., and Murphy, R. D.: Two-Dimensional Wind-"me1 9.

Tests of a Model of an F9F-5 Airplane Wing Section Using a High-

Speed Jet Blowing over the Flap; Part I - Tests of a 6-~oot Chord

Model. Aero. Rep. 845, David W. Taylor Model Basin, May 1953.

Boundary Layer Control for Various Modifications 10. Goldsmith, John: of Sweptback Wings. Rep. R-13037-5, East Hartford Research Dept., .United Aircraft coo Sept . 16, 1948.

The Supersonic Blowing Jet for Wing-Lift Augmen 11. Attinello, John S.: tation. Rep. No. DR-1706, Navy Dept. Res. Div., Oct. 1954.

12. Wallace, Richard E., and Stalter, J. L.: Systematic, Two-Dimensional Tests of an NACA 23015 Airfoil Section With a Single-Slotted Flap and Circulation Control. Aero. Rep. 120, Municipal University of Wichita, Aug. 1924.

13. Kelly, Mark W., and Tolhurst, William H., Jr.: Full-Scale Wind- Tunnel Tests of a 35O Sweptback Wing Airplane With High Velocity Blowing Over the Trailing-Edge Flaps. NACA RM A55109, 1953.

14. Allen, Julian H., and Vincenti, Walter G . : Wall Interference in a

Two-Dimensional-Flow Wind Tunnel With the Consideration of the Effect of Compressibility. NACA Rep. 782, 1944.

15. Ames, Milton B., Jr., and Sears, Richard I . : Determination of

Control-Surface Characteristics from NACA Plain-Flap and Tab Data.

NACA Rep. 721, 1 9 4 1 .

16. Kelly, John A,, and Hayter, Nora-Lee F.: Lift and Pitching Moment

at Low Speeds of the NACA 64A010 Airfoil Section Equipped With Various Combinations of a Leading-Edge Slat, Leading-Edge Flap, Split Flap, and Double-Slotted Flap. NACA TN 3007, 1953.

17. Wenzinger, Carl J., and Harris, Thomas A . : Wind-Tunnel Investigation of an NACA 23012 Airfoil With Various Arrangements of Slotted Flaps. NACA Rep. 664, 1939.

18. Williams, J.: An Analysis of Aerodynamic Data on Blowing Over Trail

ing Edge Flaps for Increasing Lift. Rep. No. 17,027, British

A.R.C. Performance Sub-Committee, Sept. 6, 1954.

19. Ames Research Staff: Equations, Tables, and Charts for Compressible Flow. NACA Rep. 1135, 1953 Figure 1.- The h o r i z o n t a l d i v i d e r s i n s t a l l e d i n the 7- by IO-foot wkd tunnel to provide a 4- by 10 foot test section; view downstream.

Figure 2.- The model installed i n the &- by 10-foot t e s t section.

Figure 3.- A detailed v i e w of the m o d e l with flap A showing the exit, of the nozzle.

chombar I " X to"

I FLAP COORDINATES 1

r = Lower U p p e r r = 0.77 -3.30 0.69 -5.63

2.62 - 1.75 1.39 -5.87

- 1 1 7 2.08 -5.96 : 7 8

5.17 -0-27 P 7 l 1 i i 3 1

10.27 I .74 17.70 -5.20 15.30 3.19 33.33 -4.38 20.34 4.05 66.65 -2.42 X i

25.34 4,46 83.35 - 1.34

= 0.6541 30.34 4.51 100.00 -0.21 = 0.6555 33.33 4.38 66.65 2.42 83.35 1.34 100.00 0.21

f =- 0.09 2 I

r = 0 . 1 0 4 ~ 1 Detail A

Figure 4.- The NACA 0006 airfoil showing the 30-percent-chord flap A, the 13-percent-chord

leading-edge flap, and the nozzle details.

TRemovable fairing used to form the single-slotted flap

{?.----p +~Ff!

- - - X I ~ ~ 0 . 7 5 (note: r a d i i of f l a p s -$=0.75 B,D,E, & F are tangent FLAP D SINGLE -SLOTTED FLAP ( w i t h fairing 1 to airfoil or FLAP A

~ ~ ~ < r ~ T , , , , , - A i r f oi I chord I i ne

- Filler b l o c k - 7 I X ~ ~ 0 . 7 5 Adjustable plate--/ +=0.75 FLAP B FLAP E Contour of

- Faired tongent to circular

contour of F l a p B contour of F l a p B arc ond - point - $0.85 +h0.75 F L A P F F L A P C configurations tested.

flap

Figure 5.- The various

w W . - ... . . ..

.. ..

\ Figure 6.- Sketch of flaps A, B, and C deflected 6 0 ' .

_ _ ~- . -. . ... .. . . . .

I The symbols indicote the position of the nose 0.70 c of the f l a p for the vorlous deflections I I 0 I 2 3 4 5 6 7 ' f 3 percent chord (a) Single-slotted f l a p .

The symbols indicote the p o s i t i o n of the n o s e of the f l o p for the v a r i o u s deflections o e x t e n d e d p o s i t i o n 0 i n t e r m e d i a t e p o s i t i o n 0 o g o i n s t t h e n o z z l e o - ! o

I I I I -n

4 1

Flap in extended position, 6 560' 0 I 2 3 4 5 X f , percent chord Flap against nozzle, 8 -60" ( b ) Flap A.

Figure 7.- The selected locations of the nose of the s i n g l e - s l o t t e d f l a p and of f l a p A f o r various f l a p deflections.

4 1 5 . 6 5.2 4.8 4.4 4.0 3 . 6

-

+ ” 3.2 r W

.-

u . 2 . 8

‘c Q)

” 2.4

.I- -9-

. -

2.0 c .- c 0 1 . 6 Q) cn 1.2 .8 . 4 - . 4 7 8 -1.2 -2 Section angle of attack, Q , deg Figure 8.- Effect of nose-flap d e f l e c t i o n on t h e 1 i f . t of t h e model w i t h t h e s i n g l e - s l o t t e d flap d e f l e c t e d 50°; R = 4.0X106.

4 2 5.6 Fairing .5.2 4.8 4 . 4 4.0 3 . 6 . . , 3.2 c W

.-

E 2.8 + W 2.4 c Lc - .

2 .o c .- t 0 1.6 W m 1 . 2 .8 -4 - . 4 7 8 - 1 . 2 S e c t i o n a n g l e of a t t a c k , a , d e g Figure 9.- Effect of slotted-flap deflection on %he lift of the model with the nose flap deflected Oo and 30'; R = 4.0XlO".

4 3 c\ -D 4.Q 3.6 3.2 2 . 8 I i

2.0 ---- r------i

1 . 6 -- r- I-

I 1.2 --I

.8 - -7- 7

0 0 4.0 I 0 500 -_- _ _ _ 0 0 4.0

0 600 __- _ _ _

0 0 4.0 1.554 1.1 6 0.0082 0.1 I 7 3.3 'I Q 0" 1.680 1.20 0.0129 0.273 2.3 2.3 0 1.518 1.16 .0104 .217 ,0129 .272 2.3, .0129 .273 2.3 A 1.680 1.20 IQ' 60" 1.683 1.23 .0128 .277 2.3 -1.2 - 1 6 -12 -8 - 4 0 4 8 12 1 6 20 -24 -20 - 1 6 -12 -8 - 4 0 4 8 1 2 1 6 20 Section angle of attack, a , deg Figure 10.- Effect of blowing on the lift Figure 11.- Effect of trailing-edge flap deflection on the lift of the model of the model with flap A undeflected; S/C = 0.001~0; 6 , = 00. with flap A in the extended position with and without blowing; s/c = 0.00110; 8 , = 0 ' .

5.6 5.2 4 . 8 4 . 4 4 . 0 3 . 6

-

2 3 . 2 c W .- V GI 2 . 8 w- W 2.4 t c - .

2 .o

c .- t 0 1.6 W v, 1 . 2 .8 . 4

-

R X l C i 6 - . 4 4.0 4.0 3.3 -8 2.3 2.3 2.3 - 1 . 2 S e c t i o n a n g l e of a t t a c k , C Y , deg (a) Variable a.

Figure 12.- Effect of blowing on t h e lift of t h e model with f l a p A i n t h e extended position; s / c = 0.00110; 6 = 30'; 6, = 0 .

0 .04 .08 . I 2 . I 6 .20 . 2 4 .28 Section mass-flow coefficient, eo Section jet-momentum coefficient, cp ( b ) Variable n o z z l e f l o w .

Figure 12.- Concluded.

I 5 . 6 5 . 2 4.8 4 . 4 4 . 0 3 . 6

-

I + 3 . 2 c W .- tz 2.8 + W 2.4 t +

. -

2.0 c .- t 0 1 . 6 W

cn

1.2

.e

. 4 C R CP X l c i 6 - . 4 --- 25" _ _ _ 0 0 4 . 0 --- 30" -_- 0 0 4.0

--- 0 350 _ _ _ 0 0 4 . 0

0.0068 0.085 4.0 -.E 1.685 1 . 1 7 Q 25" .083 4 . 0 .0068 1.617 1 . 1 6 q 3 0 : .080 4 . 0 .0067 1 . 1 7 1.584 Q 35 -1.2

S e c t i o n a n g l e of a t t a c k , Q , d e g

Figure 13.- Effect of nose-flap d e f l e c t i o n on t h e l i f t of t h e model with f l a p A i n t h e extended p o s i t i o n with and without blowing; s/c = 0.00110.

5 . 6 4 . 8 4 . 4 4 . 0 3 . 6

-

2 3.2 c W .

C 2 . 8 + Q,

" 2.4

.c - .

2 . 0 c .- 0 1 . 6 W m 1 . 2 .8 . 4 R Mi6 - . 4 --- 0 00 - - _ 0 0 4.0 --- 350 - _ _ 0 0 4.0

0 400 _ _ _ _ _ _ 0 0 4.0

7 8 b 0" 1.683 1.23 0.0128 0.27 7 2.3 b 35'1.671 1.19 .0127 0.270 2.3 b 40" 1.661 1.22 .0124 ,267 2.3 ~~ -1.2 -24 -20 - 1 6 - 1 2 -8 -4 0 4 8 1 2 1 6 20 24 28 Section angle o f attack, a , deg (b) 6 = 60° Figure 13.- Concluded.

5.6 5 . 2 4.8 I \ 4 . 4 4.0 . 3 . 6

-

t” 3.2 c Q) .- r;I 2.8 + W 2.4 t + . -

2 .o

c .- t 0 1 . 6 Q)

cn

1 . 2 .8 -4

-

- - - . 4 R CP X l C P 7 8 4.0 --- --- 0 1.19 0.0083 0.1 I 6 3.3 1.537 .265 2.3 0 1.646 1.22 .0126 -1.2 -20 -I 6 - 1 2 -8 -4 0 4 8 1 2 1 6 20 24 28 -24 S e c t i o n a n g l e o f a t t a c k , a , d e g

Figure 14.- Effect of blowing on the lift of the model with flap A

undeflected; s / c = 0.00110; En = 35 .

I

4.0 0 1.380 1 . 1 6 0.0057 0.060 4.0 0 0 0 1.560 1.19 ,0084 . I 19 3.3 A 1.450 1.16 ,0104 ,197 2.3 b 1.635 1.18 ,0124 .265 2.3.

- - - I -16 -12 -8 -4 0 4 8 1 2 1 6 20 24 28 -16 -12 -8 -4 0 4 8 1 2 1 6 20 24 Section angle of attack, Q , deg

Figure 15.- Effect of blowing on t h e l i f t

Figure 16.- Effect of blowing on the l i f t

of t h e model with f l a p A deflected 2 0 ' of the model with f l a p A deflected 35'

i n t h e extended position; s / c = 0.00110; i n t h e extended position; s/c = 0.00110;

6, = 35

6, = 35'.

5 . 6 5.2 4.8 4 . 4 4.0 3 . 6 . . .

+ ” 3.2 c a J .- V 2.8 Y a J 2.4 c +

. -

2.0 .- c 0 1.6 8-0”; 8,=35; 0 Cp.0 a, A

cn

I 1.2

.a

. 4 C

- -

Po Tstd - . 4 --- --- 0 0.0055 0.055 4.0 1.374 1.19 .097 3.3 0 1.450 1.12 .0075 2.3 A 1.507 1.08 .0090 .I50 -.E .0106 ,192 2.3 h 1.464 1.18 .0126 ,270 2.3 0 1 . 6 8 0 1.2 I .~ -1.2

-

c -20 -16 -12 -8 -4 0 4 8 1 2 1 6 2 0 24 28 Section angle o f attack, Q , deg (a) Variable a.

Figure 17.- Effect of blowing on the l i f t of the model with f l a p A deflected 50° i n t h e extended position; S/C = 0.00110; 6, = 33’.

51 *

0 .04 .08 . I 2 .I6 .20 .24 .28 Section mass-flow coefficient, cq Section jet-momentum coeff icicnt, c,, (b) Variable nozzle f l o w .

Figure 17.- Concluded.

5 . 6 5 . 2 4.8 4.4

/'

4.0 3 . 6

-

V "

- 3 . 2

c W .- V G 2.8 + W 2.4 t +

. -

I

2 .o

c /" .- c 0 1.6 a,

c n

1 . 2 / / .8 .4 --- 0

--- * 0

0 4.0 o 1.238 1.09 0.0045 .037 4.0 1.1 7 .0067 .077 4.0 0 1.539 - . 4 1.14 .0076 .lo1 3.3 A 1.466 b 1.573 1-18 -0083 .I 19 3.3 a 1.303 1.10 .0084 .I38 2.3 .I65 2.3

-.a 0 1.391 1.10 -0095

.I99 2.3 0 1.473 1.16 .0106 1.19 .0127 2 7 0 2.3 n 1.671 -1.2

-

% -20 -16 -12 -8 -4 0 4 8 1 2 1 6 20 24 28 S e c t i o n a n g l e o f attack, a , deg (a) Variable a.

Figure 18.- Effect of blowing on t h e l i f t of t h e model with f l a p A

deflected 60° i n t h e extended position; s / c = 0.00110; 6n = 35 .

5 3

------

-----

-----

---------- -------------

0 .002 ,004 .006 .008 .OlO .012 .014 0 .04 .08 .12 .I6 .20 .24 .28 Section mass-flow coefficient, cq Section jet-momentum coefficient, cp ( b ) Variable n o z z l e flow.

Figure 18.- Concluded.

5.6 5.2 4.8

"b

4 . 4 4 . 0 3.6 , . .

-- 3.2

c W .

rc 2.8

Lc W

* 2.4

c 'c -

x

2 .o

c .- c

F

u 1.6 0)

cn

' : 1

1.2 / - s=o0; s,=35: c ; o

I/?

.8

;//I

. 4 R p'i - C Q CP X K i 6 P o - - - 0 0 0 4.0 o 1.208 1.07 0.0032 0.028 4.0 1.08 .0050 .043 4.0 - 0 1.289 - . 4 A 1.463 1.1 I .0062 .066 4.0 b 1.595 1.1 9 ,0067 .080 4.0 ,0085 .I22 3.3 1.19 . n 1.584 - . 8 o 1.420 1.21 .01 00 .I84 2.3 0 1.580 1.2 I .01 17 .240 2.3 n 1.665 1.17 .0126 .270 2.3 -1.2 ~ -24 -20 -16 -12 -8 -4 0 4 8 1 2 1 6 20 24 28 Section angle of attack, CY, deg

Figure 19.- Effect of blowing on t h e l i f t of t h e model with f l a p A

deflected 65' i n t h e extended position; s / c = 0.00110; S, = 35'.

5.6 5 . 2 4.8 4 . 4 4.0 3 . 6 “I V ,.

- 3.2

c W .- C 2.8 + W 2.4 t + - .

2.0 c .- c 0 1 . 6 a , s=o0; s,=35; cp=o

cn

1.2 / / / ~ .8 R / p,j Ttj C 9 CP X K ?

_-

,’ 7 Tstd

-4 - - - --- 0 0 4.0 I 0.0018 0.004 4.0 [7 1.028 1.08 A 0 1.066 1.06 .0024 . O l O 4.0 0 / .0028 .01 5 4.0 A 1 . 1 0 3 1.08 / .0033 .020 4.0 b 1.1 37 1.07 .0041 .030 4.0 0 1.2 I 3 1.08 __ - . 4 .0048 .040 4.0 o 1.302 1 . 1 I .0054 .050 4.0 0 1.41 I 1.1 2 0061 .066 4.0 n 1.5 I 9 1.1 5 7 8 .0074 .I 18 2.3 v 1.253 1.10 .0092 .I60 2.3 v 1.407 1.1 2 - 1 . 2

S e c t i o n angle of attack, Q , d e g

(a) Variable a.

Figure 20.- Effect of blowing on the l i f t of t h e model with f l a p A deflected 5 0 ’ in t h e position against t h e nozzle; s/c = 0.00110; 6, = 350.

5 6

.002 .004 .006 ,008 .OlO .012 .014 0 .04 .08 .I 2 .I6 .20 .24 .28 0

Section mass-flow coefficient, cq Sect ion jet- momentum coefficient, cB ( b ) Variable nozzle flow.

Figure 20.- Concluded.

3.6

-

2 3.2

c .

G 2.8

.c Q) 2.4 -I- + - .

2 .o

c

.-

-I 0 1.6 Q)

cn

t r

1.2 R .8 CQ CP XIo-(

--- - - _

0 0 4.0 .4 1'; 1.027 1.01 0.0019 0.005 4.0 1.06 I 1.03 .0024 . O l O 4.0 A 1.097 1.01 .0030 .01 5 4.0 h 1.113 1.07 .0031 .018 4.0 n 1.17 I 1.1 2 .0039 .027 4.0 0 1.276 1.1 3 .0051 .044 4.0 0 1.441 1.1 7 .0057 .060 4.0 -.4 0 1.190 1.06 .0058 .083 2.3 v 1.240 1.07 ,0077 .I05 2.3 V 1.392 1 . 1 I .0096 .I63 2.3 0 1.507 1.1 5 . O IO6 .204 2.3 0 1.600 1.17 .01 I 5 .240 2.3 0 1.652 1.18 .OI 18 .264 2.3 ~~ - I 7

..-

4 8 1 2 1 6 20 24 28 -24 -20 -16 -12 -8 -4 0

S e c t i o n a n g l e of a t t a c k , Q , d e g

(a) Variable a.

Figure 21.- Effect of blowing on t h e l i f t of t h e model with f l a p A d e f l e c t e d 60° i n t h e p o s i t i o n a g a i n s t t h e nozzle; s/c = 0.00110;

.002 ,004 .006 .008 .010 .012 ,014 0 .04 .08 .I 2 .I6 .20 .24 .28 0 Sect ion j e t -momentum coefficient, cp Section mass-flow coefficient, cq (b) Variable nozzle flow.

Figure 21.- Concluded.

_.._ .. .. . . . . .

5.6 5.2 4.8 4.4 4.0 3.6

-

..

+ 3.2 c W .- C 2.8 + W 2.4 c v-

. -

2.0 t .- c 0 1.6 W A

c n

I \ 1.2

1.- I

.8 R

' Ptj Ttj

- / - C a CB XlC6 / ' Tstd -4 0 0 4.0 --- --- 0 o 1.062 1.03 0.0029 0.010 4.0 .021 0 1.135 1.04 .0034 4.0 A 1.200 1 . 0 6 .0041 .030 4.0 b 1 . 2 7 I 1.09 .0046 .039 4.0 o 1.386 1 . 0 7 .0055 .054 4.0 -.4 0 1.200 1.03 .0069 .085 2.3 0 1.235 1.1 I ,0072 ,099 2.3 n 1.353 1.1 4 .0090 .I5 I 2.3 v 1.507 1.18 ,0105 .209 2.3 7 8 V 1.572 1.19 .01 I2 .235 2.3 o 1 . 6 4 1 1.20 .01 18 .247 2.3 - I -1.2 Section angle o f attack, Q , d e g (a) Variable a , .

Figure 22.- Effect of blowing on the lift of the model with flap A deflected 70' in the position against the nozzle; s / c = 0.00110; 8, = 350.

6.0 5.0 1.0 .

c u

cn

-1 .o

0 .002 ,004 .006 .008 .010 .012 ,014 0 .04 .08 .I 2 .I6 .20 .24 .28 Section mass-flow coefficient, cq Section j e t-momentum coefficient, cc (b) Variable nozzle flow.

Figure 22.- Concluded.

_-_-----

5.2 4 . 8 4 . 4 4 . 0 3 . 6 . .

+ 3.2 c Q) .- V G 2.8 + a, .

2.4 c w . -

2 .o

c

0 2

.- c 0 1 . 6 a ,

cn

__ 1 . 2 .8 . 4 R CP X l C T 6 0 4.0 0 4.0 0 1.441 I. I 4 0.0057 0.060 4.0 - . 4 d 1.593 1 . 1 2 .0053 .060 4.0 0 1.658 1.20 .0124 2 4 6 2.3 .0104 .242 2.3 6 1.794 1.1 6 Flagged symbols indicate wing-flop slot is sealed - ~ -1.2

4 a 1 6 24 28

-24 -20 -16 - 1 2 -8 -4 0 1 2 20

Section angle of attack, Q , deg

(a) Variable a , .

Figure 23.- Effect of sealing the wing-flap slot on the lift of the

model with flap A deflected 60' in the position against the

nozzle; s / c = O.OOUO; 6n = 35'.

I

5.6 5.2 4.8 4.4 4.0 3 . 6

-

o ..

-I- 3 . 2 r W .

rt 2.8 Y- W 2.4 -I-

-

. -

2 .o

c .- -I o 1.6 W

m

1.2 -8 R pt, Tt,

- -

C P C P X 1 c i 6 Po Tstd -4 - - - 0 --- 0 0 4.0 1.069 1.02 0.0014 0.006 4.0 0 1 . 1 31 1.03 .001 8 . O I I 4.0 A 1.1 99 1.04 ,0022 . O I 6 4.0 h 1.264 1.04 .0031 .02 I 4.0 n 1.385 1.10 -0031 .030 4.0 -.4 0 1 . 1 64 1.06 .0036 0 4 2 2.3 0 1.232 1.08 .0040 .054 2.3 n 1.336 1 . 1 I ,0050 .081 2.3 7 8 v 1.417 1 . 1 1 .0056 .096 2.3 C ' 1.503 1.1 2 .0061 .I 16 2.3 o 1.782 1 . 1 6 .0075 .I69 2.3 -1.2 S e c t i o n a n g l e of a t t a c k , a , d e g (a) Variable a.

Figure 24.- Effect of blowing on the lift of the model with flap A

deflected 50' in the position against the nozzle; s / c = 0.00065; 6 , = 350.

6 3 c V .)

c c a# .

.-

no r a# V c r

.- -

C

.-

c a# m Section mass-flow coefficient, cq Section jet-momentum coefficient, cp ( b ) Variable nozzle flow.

Figure 23.- Concluded.

. .”

0 .04 .08 .I 2 .I6 .20 .24 .28

0 .002 .004 .006 .008 .OlO .012 ,014

Sect ion j e t - momentum coefficient, cc

Section mass-flow coefficient, cQ (b) Variable nozzle flow.

Figure 24. - Concluded.

4.4 4.0 3 . 6 3.2 2 . 8 . . .

," 2.4

c a3

.-

V E 2.0 Y- W 0 1 . 6 c Y- - .

1.2 E

.-

c 0 .8 0) v) .4

0 _ _ - _ _ _

0 0 4.0 0 1.070 1.0 I 0.0014 0.006 4.0 0 1 . 1 41 1.0 I ,0020 . O I 2 4.0 --- --- 0 0 0 4.0 A 1.190 1.06 .0023 . O I 6 4.0 0 1.1 12 1.04 0.0009 0.005 4.0 h 1.244 1.08 .0026 .02 I 4.0 0 1.220 1.06 .0012 . O l O 4.0 n 1.381 1 . 1 I .0032 .031 4.0 A 1.333 1.06 .0016 0 1 4 4.0 0 1.5 19 1 . 1 I .0036 .040 4.0 h 1.441 1.10 .0017 .01 0 4.0 0 1.246 1.06 ,0044 .061 2.3 n 1.678 1.10 ,0023 0 2 0 4.0 4 1.340 1.04 .0052 .082 2.3 0 1.301 1.05 -0026 .040 2.3 V 1.427 1 . 1 5 .0056 . I O 1 2.3 0 1.476 1.1 I .0032 .060 2.3 1 7 1.522 1 . 1 I ,0062 .I2 I 2.3 n 1.63 I 1.1 I .0037 ,078 2.3 0 1.728 1.14) .0073 . I 6 1 2.3 v 1.825 1.1 I . I O 2 2.3 ,0044 -20 -16 -12 -8 -4 0 4 8 1 2 1 6 2 0 -20 -16 -12 -8 -4 0 4 8 1 2 1 6 20 Section angle of attack, a , deg Figure 25.- Effect of blowing on t h e l i f t Figure 26.- Effect of blowing on t h e l i f t

of the model with f l a p A deflected 6 0 ' of t h e model with f l a p A deflected 50'

i n the position against t h e nozzle; i n the position against t h e nozzle; s/c = 0.00065; 6, = 35'. s/c = 0.00036; tjn = 35'.

3 . 6 3.2 2.8 0 " 2.4 ..

c c

2 2 . 0

u

.-

y.

+ 1.6 c *- 1.2

-

C

2 .8

c u a l 0 8 . 0 0 4.0 R 1.249 1.05 0.0014 0.0 I I 4.0 0 1.504 1.08 .0020 .02 I 4.0 C a Cp A 1.628 1.08 .0022 .027 4.0 - . 4

--- ' h 1.257 1.04 .0024 .033 2.3 _ _ -

0 0 4.0

c ' n 1.293 1.02 .0025 .036 2.3

1.556 1.02 0.0008 0.010 4.0 0 1.325 1.08 .0026 .043 2.3 0 1.443 1.01 .0012 ,02 I 2.3 7 8 ' 0 1.469 1.10 .0033 .060 2.3 A 1.539 1.03 .0014 .026 2.3 n 1.628 1 . 1 2 .0038 ,080 2.3 h 1.656 1.01 .0016 .032 2.3 v 1.818 1.1 2 ,0044 . I O 0 2.3

I I I n I 8 0 7 1.04 .0019 .041 2.3

-1.2 -20 -16 -12 -8 -4 0 4 8 12 1 6 -20 -16 -12 -8 -4 0 4 8 12 1 6

Section angle o f attack, a , deg

Figure 28.- Effect of blowing on the l i f t

Figure 27.- Effect 'of blowing on the l i f t

of the model with f l a p A deflected 50' of the model with f l a p A deflected 60° i n the position against the nozzle; i n t h e position against t h e nozzle; S/C = 0.00017; 6, = 3 3 O .

S/C = 0.00036; 6, = 35'.

4.0 3 . 6 3.2 2.8 2.4

c

c C 2 . 0 W .

u

.-

cc Y- W cc

.-

c

.-

4- V W v) ptj Ttj R

- -

ca CP

c. . ,r 1.599 1.05 0.0005 0.01 0 4.0

- . 4 -=-

Po Tstd 1.870 1.05 .0012 .01 5 4.0 - - - --- 0 0 0 4.0

f' A 1.428 1.03 .0012 .020 2.3

o 1.855 1.00 0.0076 0.0 I 7 2.3

- h 1.529 1.03 .0014 .025 2.3

0 1.848 1.02 ,0133 ,029 2.3 0 1.605 1.06 .OO I 6 -029 2.3 - U T * - A 1.821 1.03 .0150 .032 2.3 0 1.805 1.06 .0020 .040 2.3 b 1.798 1.03 .0199 .040 2.3 -1.2 . - - - -

----

-20 -16 -12 -8 -4 0 4 8 1 2 1 6 20 -20 -16 -12 -8 -4 0 4 8 1 2 1 6 20

Section angle o f attack, 0 , deg

Figure 29.- Effect of blowing on the l i f t

Figure 30.- Effect of spanwise extent of

of t h e model with f l a p A deflected 6 0 '

blowing on t h e l i f t of t h e model with i n t h e position against t h e nozzle; f l a p A-against the nozzle; s / c = 0.0017; s / c = 0.00017; 6, = 35'.

6 = 60°; 6 , = 3 5 O .

B

Figure 32.- Effect of blowing on the l i f t

Figure 31.- Effect of trailing-edge

of the model with f l a p B deflected 50'; f l a p deflection on the l i f t of the s/c = 0.00110; 6, = 35O.

model with f l a p B; s/c = 0.00110; 6, = 3 5 O .

5.6 5.2 4.8 s=o”; s,=35: cp=o

’ p t j Ttj R

/ - - CO CP XlO?

Po Tstd 0 0 4.0 0 1.062 1.05 0.0024 0.010 4.0 3 1.046 1.07 .0028 .OI 5 4.0 h 1.1 27 1.05 .0033 .020 4.0 h 1.161 1.06 .0038 .026 4.0 D 1.203 I. I I .0041 .032 4.0 3 1.274 1.09 .0048 .041 4.0 0 1.387 1.1 4 ,0060 .060 4.0 2 1.232 1.09 .0079 .I06 2.3 7 1.387 1 . 1 0 .0097 .I61 2.3 7 ’ 1.5 I I 1 . 1 5 .01 14 .223 2.3 -24 -20 -16 -12 -8 -4 0 4 8 1 2 1 6 2 0 24 28

Section angle of attack, Q , deg

(a) Variable a.

Figure 33.- Effect of blowing on the lift of the model with flap B

deflected 60°; s/c = 0.00110; 6, = 3 5 O .

0 .04 .08 .I 2 .I6 .20 .24 .28

Sect ion j e t -momentum coefficient, cp Section mass-flow coefficient, cq (b) Variable n o z z l e flow.

Figure 33.- Concluded.

4.4 4.0 3 . 6 3.2 2.8 2.4 2.0 1 . 6 1.2 . 8 .4 I

_ _ _ _ _ _

0 4.0

i 1.062 1.03 0.0026 0.0 I I 4.0

1 '

- 0 1.1 32 1.05 .0034 .021 4.0

/ A 1.267 1.08 .0048 .041 4.0 / h 1.360 1.09 ,0057 .053 4.0 0 1.404 1.08 .0060 .058 4.0 -.4 0 1.445 1.08 .0062 ,064 4.0 0 1.171 1.05 .0069 .082 2.3 n 1.230 1.05 .0078 .I03 2 . 3 : 8 -2 v 1.358 1.1 0 .

.0097 .I 6 0 2.3 v 1.492 1.14 210 2.3 .0112 0 1.625 1 . 1 5 .0124 ,265 2.3

-- ---- -1.2

I -20 -16 -12 -8 -4 0 4 8 12 1 6 2 0 24 28 -16 -12 -8 -4 0 4 8 1 2 1 6 20 Section angle o f attack, a , deg

Figure 34.- Effect of blowing on the l i f t

Figure 35.- Effect of trailing-edge f l a p

of t h e model with f l a p B deflected 70';

deflection on t h e l i f t of t h e model

s/c = 0.00110; 8, = 3 5 O .

with f l a p C; s / c = 0,00110; 6, = 3 5 O .

I

5 . 6 5 . 2 4.8 4 . 4 4.0 3 . 6

-

,.

+ 3 . 2 c Q)

.-

rt: 2.8 + Q) 2.4 c .c

. -

2.0

! =

.-

t 0 1.6 Q)

m

1 . 2

A

.8

L -

-4 --- --- 0 0 0 4.0

2 I

1.032 1.01 I 0.0019 0.006 4.0

-

0 1.064 1.01 .0026 .OI I 4.0 7- A 1.106 1.04 .0032 .018 4.0 /

/

b 1.121 1.07 .0034 ,019 4.0

--

0 1.275 1.09 .0049 - . 4 .042 4.0 0 1.435 1.1 2 ,0061 .064 4.0 0 1.221 1.05 .0078 . I O 1 2.3 0 1.381 1.10 .0100 .I64 2.3 7 8 v 1.511 1.14 .0113 .213 2.3 V 1.654 1 . 1 5 .0126 .285 2.3 - 1 . 2

-24 -20 -16 - 1 2 -8 -4 0 4 8 1 2 1 6 20 24 28

Section angle of attack, a , deg

Figure 36.- Effect of blowing on t h e l i f t of t h e model with f l a p C

d e f l e c t e d 50'; s / c = 0.00110; tjn = 3 5 O .

7 3 c 5.6 I I I I I 5.2 4.8 4.4 4.0 3 . 6 # . , +" 3.2 e Q)

.-

V 2 . 8 + Q) 2.4 w

. -

R Tt I

-

C a clr X l b ' Tstd

_ _ _

0 0 4.0 1.03 0.0036 0.020 4.0 1.02 .0038 -024 4.0 1.07 .0042 .030 4.0 1.1 0 .0050 ,043 4.0 -.4 1.1 2 ,0062 ,063 4.0 1.09 ,0078 . I O 3 2.3 1.1 0 0096 -158 2.3 1.1 5 .OI 18 ,240 2.3 1.1 5 .0124 .270 2.3

_. i u

-1.2 1 2 1 6 20 24 28 -24 -20 -16 - 1 2 -8 -4 0 4 8

S e c t i o n angle o f attack, Q , deg

(a) Variable a.

Figure 37.- Effect of blowing on t h e l i f t of t h e model with f l a p C

deflected 60°; s/c = 0.00110; 6 , = 35'.

0 .002 .004 .006 .008 .OlO ,012 .014 0 .04 .08 . I 2 .I6 .20 .24 .28 Section mass-flow coefficient, cq Sect ion j e t -momentum coef f icicnt, cc (b) Variable nozzle flow.

Figure 3 7 . - Concluded.

5.6 5.2 4.8 . 4.4 4 . 0 3 . 6

-

+ ’ 3 . 2 c W

.-

Z 2 . 8 Y- W 2.4 c c - .

b \ 2 .o c .- c 0 1.6 W /\ 8-0”; 8,=35: cp=O cn c ‘ \ 1 . 2 . 8 R C a CP xrb6 . 4 0 0 4.0 1.126 1.07 0.0034 0,020 4.0 0 1.202 1.08 .0042 .032 4.0 A 1.229 1.09 ,0045 ,036 4.0 b 1.284 1.06 .0052 .043 4.0 n 1.347 I. 14 ,0055 .053 4.0 - . 4 0 1.437 1.14 .0061 .065 4.0 0 1.215 1 . 1 2 .0074 . I O 0 2.3 n 1.345 1 . 1 I .0096 . I % 2.3 - . 8 v 1.536 1.1 5 .O I 14 .222 2.3 V 1.637 1 . 1 8 .0122 ,269 2.3 -1.2 Section angle of attack, O , deg Figure 38.- Effect of blowing on the l i f t of the model with flap C deflected 70’; S / C = 0.00110; 6 , = 33’.

... ._ . ..

I

5.6 5 2

I L

4.8

I I

4.4

I

4.0

I

3.6

-

3.2 c Q) .- G 2.8 Y a 2.4 t v-

. -

2 .o c .- c 0 1 . 6 Q) m

m

1.2 I

LL /'

.8 R Tti

-

/ - CQ CP X l C i 6 .4 Tstd / Po / 0 _ - - _ _ _ 0 4.0 / 1.076 1.03 0.0024 0.0 I I 4.0 . / I 0 1.096 1 . 0 I .0026 . O I 3 4.0 / A 1.1 30 1.01 .OO'L9 . O 18 4.0 / h 1.1 67 1.01 -0033 -022 4.0 ,/ n 1.33 I 1.09 ,0044 .041 4.0 -.4 0 1.532 I. I I ,0055 .062 4.0 0 1.279 1.08 .0071 .IO5 2.3 n 1.445 1 . 1 I .0088 .I61 2.3 7 8 v 1.593 1.14 .0099 201 2.3 1.738 1.15 .0113 . 2 5 4 2 . 3 I -1.2 ~ -24 -20 -16 -12 -8 -4 0 4 8 1 2 1 6 20 24 28 Section angle of attack, Q , deg (a) Variable a.

Figure 39.- Effect of blowing on the lift of the model with flap C S, = 0'; s / c = 0.00110.

deflected 50' and 7 7 I 6.0 Qu 5.0 (deg) X I O ’ ~ I 2.3

6 E 4.0

I

0 -17 2.3 . ---u -

4.0 1

‘9 -17 4.0 1 I

l - T l -

3.0 2.0

r

6 .04 .08 .I 2 .I6 .20 . 2 4 .28

Section mass-flow coefficient, cq Section jet-momentum coefficient, cp ( b ) Variable nozzle f l o w .

Figure 39.- Concluded.

5 . 6

Tr

5.2

Tr

4.8

Tr

4 . 4 l / d . k .

4.0

7-E

3 . 6

-

'E

. .

c 3.2 r: W .- rZ 2.8 cc W 2 . 4 t cc

7 - 1 7

- .-

2.0 c

kpkp"

.- c 0 1 . 6 W m

PIp

1.2

/I I

-8

/I

-4 R Tt I

-

C a CP x 1 p Tstd --- 0 4.0 1.07 0.0042 0.031 4.0 1.06 .0051 .042 4.0 - . 4 1.1 27 1.05 .0058 .060 2.3 1.230 1.08 .0080 . I O 8 2.3 'q 1.354 1.1 0 .0096 ,155 2.3 -.8 1.440 1.1 3 .0106 .I88 2.3 1 . 1 5 .0126 270 2.3 -1.2 Section angle of attack, Q , deg Figure 40.- Effect of blowing on the lift of the model with flap D deflected 50'; S / C = 0.00110; 6 , = 35'.

7 9 5.6 5.2 4.8 4 . 4 4 . 0 I 3 . 6 L V ..

+ 3.2 c W .- d- 2.8 cc W 2.4 w

. -

d?, 2.0 \ r .- t 0 1 . 6 W v, 1 . 2

J$

. 8

P

-4 R

- i ?

CCL X t c P Po Tstd - . 4 --- --- 0 0 4.0 0 1.1 26 1 . 0 6 0.0056 0,060 2.3 0 1.185 1 . 0 7 .0073 ,090 2.3

1 s

7 8 .0079 . I O 6 A 1.233 1.08 2.3 .01 I6 .245 2.3 L 1.566 I. I7 o 1.637 1 . 1 5 .0125 ,269 2.3 - - 1 . 2 -24 -20 - 1 6 - 1 2 -8 -4 0 4 8 1 2 1 6 20 24 28 S e c t i o n a n g l e of a t t a c k , O , d e g (a) Variable a.

Figure 41.- Effect of blozing on the lift of the model with flap D

deflected 60 ; s / c = 0.00110; 6, = 35’.

6.0 5.0

- 4.0

.

c C 2 3.0 .

r w a I 2.0 1.0 .

c a J cn

-I .o

0 .002 .004 .006 ,008 .OlO .012 .014 0 .04 .08 .I 2 .I6 .20 .24 .28 Section jet-momentum coefficient, c,, Section mass-flow coefficient, cq (b) Variable nozzle flow.

Figure 4 1 . - Concluded.

co P I R C a CP X I 6 6 , 0 4.0

/ i o 1.126 1.05

0.0057 0.060 2.3

- . 4 ---

0 1.246 1.08 ._ .0080 .I 13 2.3 ' A 1.280 1. 10 ,0086 . I 2 5 2.3 b 1.363 1.08 .0096 . I 5 8 2.3 ,0104 . I 9 3 2.3

I 0 1.633 1 . 1 7 . O 1 24 ,269 2.3

0 3 5 " - - - --_ I 0 Figure 42.- Effect of blowing on the l i f t Figure 43.- Effect of trailing-edge f l a p of t h e model with f l a p D deflected 70'; deflection on the lift of t h e model s / c = 0.00110; 6n = 3 5 O . with f l a p F; s/c = 0.00110; � 5 , = 35'.

5.6

n

5.2 4 . 8 \ , , .

4 . 4 4 . 0 3 . 6

-

3.2 c W .- 2 . 8 + W 2.4 t Lc - .

2 .o c .- -I o 1.6 W v, !

i 1.2 .0 -4

_i

-.4 7 8 - 1 . 2 Section angle of attack, Q , deg Figure 44.- Effect of blowing on the lift of t h e model with f l a p F deflected 50°; s/c = 0.00110; S, = 35'.

5 . 6 5 . 2 4.8 4 . 4 L 4.0 3 . 6 n.

+- 3.2 c Q) .- V rt 2.8 + Q) 2.4 +

. -

2 .o c .- t 0 1.6 Q) c n - 1 . 2

I. i

.8 R C Q CP Xld' . 4 0 0 4.0 o 1.1 01 1.05 0.0036 0.0 I 9 4.0 0 1.1 69 1.05 .0045 -030 4.0 0 A 1.235 1.06 .0052 .041 4.0 b 1.303 1.08 .0059 -052 4.0 n 1.1 I 6 1.03 .0062 .061 2.3 0 1.1 55 1.04 .0074 .083 2.3 - . 4 0 1.194 1.04 .0082 . I O 0 2.3 0 1.306 1.1 3 .0098 ,155 2.3 v 1.455 1.1 6 .OI 13 .201 2.3 -.a P 1.515 1.17 .0127 .240 2.3 o 1.558 1.1 5 .0132 .270 2.3 0 1.624 1.1 9 - .0141 -. .300 2.3 -1.2 -24 -20 - 1 6 - 1 2 -8 -4 0 4 8 1 2 1 6 20 24 28 Section angle of attack, Q , d e g (a) Variable a.

Figure 45.- Effect of blowing on t h e l i f t of t h e model with f l a p F deflected 60'; S/C = 0.00110; 6n = 35'.

. . - .... ...

i----

d\

----- I)(

-

-

. . . .

.002 ,004 .006 .008 . O l O .012 ,014 0 .04 .08 .I 2 .I6 .20 .24 .28 0 Section mass-flow coefficient, cq Section jet -momentum coefficient, cp (b) Variable nozzle flow, Figure 43. - Concluded.

--- 0 --- 0 0 4.0 1.120 1.05 0.0037 0.022 4.0 0 1.243 1.04 .0053 .042 4.0 A 1.302 1.09 .0058 .05 I 4.0 h 1.109 1.05 .0062 .059 2.3 D 1.124 1.06 .0064 .068 2.3 .0074 .081 2.3 0 1.153 1.06 0 1.181 1.08 .0079 .095 2.3 6 1.307 1.12 .0101 . I 5 6 2.3 v 1.405 1.1 I .OI I 6 -203 2.3 V 1.524 1 . 1 6 .0128 .240 2.3 o 1.562 1-14 ,0134 .270 2.3 -. I I 4 8 1 2 1 6 2 0 24 28 * Section angle of attack, Q , deg Figure 46.- Effect of blowing on the lift of the model with flap F deflected 70'; s/c = 0.00110; 6, = 33'.

4.0

3.2

@

2.4

I .6

.a

- Q "U

.2 0 -. 2 -.4 0.6 -.8

.2 0 -. 2 -.4

Section pitching-moment co efficient, c,

(a) cp = o (b) cp 0.12

Figure 47.- Effect of blowing and of flap deflection on the pitching-moment characteristics of the model with flap A in the extended position; s/c = 0.00110; 6, = 35'.

5.6

4.8

4.0

rr

+" c 3 . 2

W .

.-

.c .c W

s 2 . 4

c

-

1.6

.-

c W d3

.8

- 8

.2 0 0.2 -.4 -.6 -.8 -1.0 -1.2

Section pitching-moment coefficient, c, .

( c ) cCI 0.27 Figure 47. - Concluded , ,

5.6

4.8

4 . 0

@

*" 3 . 2 c

a

.-

.-

w

*

a 2.4 c !e 1.6

.-

c a v)

. 8

9.8 .~

.2 0 -. 2 -.4 -.6 - . 8 - 1 . 0 - 1 . 2

Section pitching-moment coefficient, c, Figure 48.- Effect of blowing on the pitching-moment characteristics of the model with flap A deflected 60° in the extended position; S / C = 0.00110; 8 1 1 = 35O.

8 9 a t C Q) .

..I)

*

*

0 ) c

=

-

C

.-

t 0) v,

.2 0 -. 2 -.4 0.6

Section pitching-moment coefficient, c,

(a) cp = o (b) cp 0.03

Figure 49. - Effect of blowing and of flap deflection on the pitching-moment characteristics of the model with flap A against the nozzle; s/c = 0.00110; 6 , = 35'.

3 . 2

r,

," C 2 . 4

0 ) .

.-

w w- Q) 1.6 t !e C

.- 0 . 8

c Q)

cn

-.0

- . 2 -.4 - . 6 -.8 -. 2 -.4 -.6 -.8

-. 2 -.4 - . 6 - . 8

Section pitching-moment coefficient, c, (a) s/c = 0.00063 ( b ) s/c = 0.00036 ( c ) S/C = 0.00017 Figure 50.- Effect of nozzle height and of flap deflection on the pitching-moment characteristics

-

of the model with flap A against the nozzle; cp = 0.03; S, = 35'.

4.8 4 . 0 3.2 2.4 I . 6 .8

- - , I - - I------

&pq I_J

-

-- ----- I.

-.8

.2 0 -. 2 -.4 - .6 -.8 -1.0 .'5 0 -.2 -.4 -.6 -.8

Section pitching-moment coefficient, c, (a) s/c = 0.001~0 (b) S / C = 0.00065 Figure 51.- Effect of blowing and of nozzle height on the pitching-moment characteristics of the model with flap A deflected 60° against the nozzle; 6, = 35 .

4 . 0

@

+" c 3.2

Q) .

.-

w - Y Q)

2.4

.6 -.8

0 -. 2 -.4 -.6 -.8 .2 0 -. 2 -.4 - .6 -.8

Section pitching-moment coefficient, c,

(a) S/C = 0.00017

(c) S/C = 0.00036 Figure 51.- Concluded.

W w

- 1

I

I

I

I

d

e

f a

4 1

(a) C Q = 0; cP = 0; R = 4.0X106 a , d e g . chord P -4.1 76.00 2 0 (b) CQ = 0.0082; cCI = 0.117; R = 3.3a06 a , d e g chord P -4.1 76.00 4.9 F l o g g e d symbol Percent chord (c) c Q = 0.0129; cp = 0.273; R = 2 . 3 ~ 1 ~ Figure 32.- Effect of angle of attack and of blowing on the chordwise distribution of pressure of the model with flap A undeflected; s / c = 0.00110; 6 , = 0.

L

Flogged symbols indicoie lower surface

I

-4 I

I

-3 I

L

-2 I

L

- '

a

i- O r f

z @

; E l .t a l (a) C Q = 0; cp = 0; R = 4.0xloS V -7 0, L

I I 1 -

I n I n

I I I

E -6 a

I I I

I I 1

-5

I I I

I I I

-4

I 1

I I I

-3

I 1

-2

\I i

* I 1

- I

-t- =L

A F It

-i

I

4 0 =k T

I I O 20 30 3 80 90 5 0 60 Percent chord (b) C Q = 0.0126; cp = 0.265; R = 2.3~J-w Figure 53.- Effect of angle of attack and of blowing on the chordwise distribution of pressure of the model with flap A undeflected; s/c = 0.00110; fjn = 35 .

. , . ,-, ... . . .. . . . .

Percent chord (a) CQ = 0; cP = 0; R = 4.OXloS Figure 54.- Effect of angle of attack and of blowing on the chordwise distribution of pressure of the model with flap A deflected 35' in the extended position; S/C = 0.00110; 6, = 33'.

* -14

I

- I3

0 0.1

I

Flagged symbols indicate lower surfacc

I

- 1 2

t

- I I

I

I

-10 -9 Q

-c

..

c

I

c -a

e , .-

I

.- Lc c

I

g -7

I

a 8

I

2 -6 u)

u) I

I

- 5

I

I

-4

L

-3

i

I

-2

I

It

- I

I

I

-c

I 20 30 40 50 Percent chord (b) CQ = 0.0056; C V = 0.060; R = 4.OXlOS Figure 54.- Continued.

I I1 I I I I1 I1 I I 1 1 1 1 1 1 1 1 1 1 l 1 1 1 1 1 1 l 1 1 1 1 1 l 1 1 l 1 l 1 1

.

-14

- I3

0 4. I F l a g g e d symbols -12

i indicale lower surfacc

- I I -10 -9 n ..

c a J -8 .

.- + + g -7 0, 5 -6 u) u) -5 -4 -3 -2 - I ' ( I O 20 30 40 50 6 0 Percent chord (c) C Q 0.0127; CP = 0.270; R = 2.3X106 Figure 34.- Concluded.

9% -14

- I3

I7 2.0 Flagged symbols ndicaie lower surface

- 1 2

- I I -10 -9 Q ..

c

= -8 Q)

.

.- w +

g - 7

u a J 5 -6 In u) ?

- 5 -4 -3 - 2 - 1 C I 60 70 1 90 100 Percent chord (a) C Q = 0; cCL = 0 ; R = 4.0X106 Figure 55.- Effect of angle of attack and of blowing on the chordwise distribution of pressure of the model with flap A deflected 50° in the extended position; s / c = 0.00110; 6 , = 33'.

I

P - I 6 . 6 - I 8.3

I

I

I

90 I O 0 ' 0 ' I O 20 30 40 50 6 0 Percent chord ( b ) C Q = 0.0075; c p = 0.097; R 3.3XLO' Figure 35.- Continued.

-14

I I I I I

Percent

- I3

~ . d e g . chord P

~ 4 p jig 1

75.45 -19.5 -10.0 " 75.71 -23.1

- 1 2

" 76.08 -21.2 o 0.3 Flagged symbols indicate lower surface

I

- I I

I

I

-10

I

-9

I

a

-

.a

I

5 .-0

.

V I

.- rc +

I

g -7

V

I

I

E 3 -6 u)

I

u)

E

I

-5

I

I

-4

I

I

-3 / 0

d-

/

z L

-2

_L

- 1

I

C

I

*

I I O : 30 40 50 Percent chord ( e ) CQ = 0.0126; ccL = 0.274; R = 2.3XI-06 Figure 35.- Concluded.

I I I I I1 l l l 1 1 l l I Ill Ill1 1 1 1 1 1 1 1 1 l 1 1 1 1 1 1 1 1 1 l 1 1 1 1

Flagged symbols indicate lower surface -I I -10 -9 a ..

t W -8 .

.- + .r

0 " -7

W -6 u) u) - 5 -4 -3 -2 - 1

@ J ) y q L

1 : 0 I O 2 6 0 70 I O 0 Percent chord (a) CQ = 0; c p , = 0 ; R = 4 . 0 ~ l - 0 ~ Figure 56.- Effect of angle of attack and of blowing on the chordwise distribution of pressure of the model with flap A deflected 60° in the extended position; s/c = 0.00110; 6 , = 35'.

-14

4 ; "/g I

- I3

Flagged symbols

- 1 2

indicale lower surface - I I -10 -9 a

-

= - 0 0 ) .- .-

-

*

g -7

aJ f -6 In In -5 -4 -3 -2 - I I O 20 3 0 6 0 70 8 0 90 100 Percent chord (b) CQ = 0.0076; cP = 0.101; R = 3.3x1O6 Figure 56.- Continued.

I ~ I I I I I l l 1 1 1 1 1 1 1 1 1 1 l l I I l I l I l 1 1 1 1 1 1 1 1 1 1 l 1 1 1 1 l I I I

I I I I

Percent chord P 75.86 -329 -39.4 n -35.7 8 ' -34.2 76. I I -52.8 " -50.7 " -49.5 'I -44.0 76.51 -29.3 " -27.9 " -26.9 1' -23.8 77.95 - I 6.4 " - I 5 . 7 " -14.6 ' " -13.2 a c E Q) .

V .- -I- + W V ln ln n IO0 Percent chord ( e ) C Q = O.Ol27; cP = 0.276; R = 2.3X106 Figwe 56.- Concluded.

-14 R -13 CQ cp X l O + 1 0 0 4 . 0

3 0.0061 0.066 4.0 c

- 1 2

Flagged symbols indicate lower surface - I I -10 -9 a ,.

c

= -a

W .

.- v +

0" -7

u a J f -6 u) v) E -5 -4 -3 -2 - I I 3 0 40 5 0 6 0 70 8 0 90 I O 0 Percent chord (a) CQ = 0 and 0.0061 Figure 57.- Effect of blowing on the chordwise distribution of pressure of the model at a constant angle of attack (a, = -4.0') with flap deflected 5O0 in the extended position; s / c = 0.00110; 6, = 35'.

I I ' 0 - I O 20 30 40 5 0 6 0 70 80 90 I( 0 Percent chord (b) C Q = 0.0073 and 0.0084 Figure 57. - Continued.

I -.--___ -- ... ._.. .. , ..... , ._ , .. 1 -14

I I I I I I

R Percent

- I3

C Q cp xio-' Cp chord P L 0.0084 0.135 2 . 3 0.135 75,.?5 -21.62 3 0 . 0 1 2 8 0 . 2 6 7 2 . 3 ,267 -23.58

- 1 2

.I35 75.71 -23.15 d symbols Flag1 .267 " -31.17 ower surface . I 35 76,;08 - 15.64 indicale .267 -21.13 - I I -10 -9 Q c Q) -8 .

.- IC -e

g -?

$ -6

v) u)

E

-5 -4 -3 -2 - I ', I O 20 30 4 0 5 0 6 0 90 I Percent chord (e) CQ = 0.0084 and 0.0128 Figure 57.- Concluded.

...

-14 -13 Flagged symbols

- 1 2

indicoie lower surface -I I -10 -9 CL ,.

c

= -8

Q) .

.- v + g -7 -6 u) u) -5 -4 -3 -2 - I T l I I I l

'

I IO 20 30 40 6 0 70 80 I O 0 Percent chord (a) CQ = 0; cp = 0; R = 4.0Xi-06 Figure 58.- Effect of angle of attack and of blowing on the chordwise distribution of pressure of the model with flap A deflected 50' against the nozzle; s / c = 0.00110; 6 , = 33'.

-14 a , deg

- I3

n 4,2 F l a g g e d symbols

- 1 2

ndicate lower surface -I I -10 -9 Q t

= - 0

0 ) .

.- -4 +

g -7

0, 5 -6 v) v) -5 - -4 -3

pr

-2

i

i

- I i i

-

- 1

90 * I

I O 20 30 40 50 60 70 Percent chord ( b ) C Q = 0.0024; cCL = 0.010; R = 4.OX1O6 Figure 58.- Continued.

Percent chord ( c ) CQ = 0.0061; cp = 0.066; R = 4.OXLO6

Figure s. - Continued

-14

- r r m

I I I I I

a , deg Percent

- I3

a , de@ chord 0.0 75.7 I

-22.5 m

- 1 2

Flagged symbols ndicaie lower surface - I I -10 -9 a ..

c C - 8 Q) .

.- + +

g -7

u al 5 -6 v) v)

E

-5 -4 -3 - 2 - I I I O 20 30 40 50 60 Percent chord (a) C Q = 0.0092; cP = 0.160; R = 2.3>u06 Figure %.- Concluded.

-14 a , de9

- I3

4.2 Flagged symbols

- 1 2

indicate lower surface - I I -10 -9 Q ..

c 0 ) -8

.-

.- .c .c

g -7

a J 5 -6 fn fn - 5 -4 -3 -2 - 1 I ( 60 70 80 90 I O 0 Percent chord (a) CQ = 0; cp = 0; R = 4.Ox1O6 Figure 59.- Effect of angle of attack and of blowing on the chordwise distribution of pressure of the model with flap A deflected 60' against the nozzle; s/c = 0.00110; 6, = 33 .

-14

H I I I

a , deg

- I3

A -4.0 0 0.2 Flagged symbols

- 1 2

ndicote lower surface - I I -10 -9 a c

5 - 8

.

.- y.

r

g -7

u Q) 2 -6 v ) v ) L a -5 -4 -3 -2 - I I I 20 5 0 6 0 8 0 90 100 Percent chord ( b ) C Q = 0.0030; cCL = 0.015; R = 4.OXloS Figure 59.- Continued.

-14 1 1 . 1 I I.

Perceni jl.deg. chord P

- I3

I . 6 -3.9 75.10 I , I .8 0.2 2.0 4.0 I '

- 1 2

-16. I -3.9 75.33 Flagged symbols I I -15.9 0.2 indicaie lower surfact II -I 5.2 4.0 -I I -10 -9 n ..

c E : -8 Q) .

.-

.c r 0 -7 0,

2 -6

in (n .L -5 -4 - 3 -2 - I I I O 0 50 6 0 80 Percent chord ( c ) C Q = 0.0059; CP = 0.060; R = 4.0X106 Figure 59. - continued.

I 1

- 1 4 1 I I I u I I I I I I :I

Percent I II

a.deg L

o - 8 . 1

A -3.9 I -

0 0.2

- 1 2 Flagged symbols

indicate lower surface

LL

LL

a

+- LL

.

.-

= lo I

E - 7 1 1 1

W

f v) - 6 F

u) 0 - IO 20 3 0 4 0 5 0 6 0 i Percent chord (d) C Q = 0.0124; cP = 0.246; R = 2.3X106 Figure 59.- Concluded.

.24 Flop location .20 OAgoinst nozzle 3.97 2.79 0 Intermediate 5 . 0 7 2.79 OExtended 5.67 2.58 .I6 d E . 1 2 .- c .- L .08 n

-

.24 .28 3 4 5 6 x f , percent chord 6 .24 5 .20 5 OAgoinst nozzle 4.12 2.83

*

0 Intermediate 4.57 2.83

-

E 4 .I6 .

z *

0 - 0 0 0 c O 3 3 2

- L 3

-

.

c 0 .- N ? t U : 2 .08 I .04 I 0 0 .04 .08 .I2 .I6 .20 .24 .28 3 4 5 6 percent chord X f , CP Figure 60.- Effect of flap position on the lift of the model with flap A; s / c = 0.00110; sn = 350.

116 E I I .007 .008 ,002 .003 ,004 .005 Y O .OOl U

-

c v

- 1

I

+E

&

c

I

I

I

x = I

I 1

.02 03 .Q4 05 06 .07 .08 0 .01

CtL

(a) ' 6 = 50' Figure 61.- The variation of the lift coefficient at zero degrees angle of attack.withthe mass-flow and the jet-momentum coefficients for the various flaps tested; s / c = 0.00110; S, = 3 5 O .

I I .005 .O 06 .007 .008 .OOl .002 .003 .004 CQ

6l

n N v

0 .01 .o 2 .03 .O 4 .05 .O 6 .07 .08

cP

( b ) 6 = 60' Figure 61.- Continued.

F l a p 0 A 0 B

0 C

A D L3 F ( F l a p A is a gainst nozzle:

i r r

I

I l l

0 .OOl .002 .003 .004 .005 .006 .007 .008 CQ N U

“ I

I

3r

r

* E

at C. = 0. I16

I

r

‘ F

L

.o I .o 2 .03 .04 .05 .06 .O 7 .08

ctL (e) 6 = 70’ Figure 61.- Concluded.

0 O.OOll0 0 .00065 0 .00036 A .00017 .003 .O 04 .005 .006 .007 .008

0 .OOl .o 02

CQ U

1 =

+

T

T I

I I

I

T

- - I I

I

I

I

I I - .I

0 .02 .04 .06 .08 .I4 . I 6

cP

(a) Flap A against the nozzle; 6 = 50°; 6 , = 35 .

Figure 62.- "ne effect of nozzle heigh% on the variation of the increment of lift coefficient with the mass-flow and jet-momentum coefficients.

C I I

0 O.OOll0

o .00065

0 .00036

I

I

-I I I

H I I

H I

0 .OOl .o 02 .003 .004

.005 .006 .007 .008 .

CQ n N

a"

- 4

#

V D

I I I I

I I I I

I

!

I I I I

I l l 1

.02 .04 .06 .08 IO .I2 . I 4 .I6

cr,c

( b ) Flap A a g a i n s t t h e nozzle; 6 = 60°; 6, = 35'.

Figure 62.- Continued.

.O 40 .40 ( c ) Flap of reference 12; 6 = 60'; no leading-edge device.

Figure 62. - Concluded.

-1" I

.002 .004 .006 .008 .o IO .012 .014

.-

G Q n N I

.O 4 .08 . I2 .I6 .20 .24 .28

cP

Figure 63.- The effect of nose-flap deflection on the variation of the increment of lift coefficient with the mass-flow and jet-momentum coefficients; flap C; s / c = 0.00110; 6 = 50°.

( F l a p t y p e e) ( F l a p t y p e f ) R e f . 4 R e f . 4 R e f . 5 R e f . 9 R e f . 12 Figure 64.- Sketches showing the arrangement of the flap and blowing systems f o r each of the referenced investigations.

.

e h.

a"

-

I Figure 65.- The effect of the jet-momentum coefficient, cPr on the variation of the increment of lift coefficient with flap deflection; flap A extended; NACA 0006 airfoil section; s / c = 0.00110; 6 , = 35'.

~ cP 0.16 .i2 . 0 8

.a

. 0 4 .02 A CinnIn clnttod flnn fin Q

50 60 70 80 50 60 70

(c) S/C = 0.00036 (d) S/C = 0.00017 Figure 66.- The effect of the jet-momentum coefficient, c on the varia P?

tion OS the increment of lift coefficient with flap deflection; flap A against nozzle; NACA 0006 airfoil section; 6n, = 3 5 O .

lotted flap,fiq.S (a) Flap B. 8 , = S O " (b) Flap C.

0 Ref.16 Double slotted flap, Cf =0.306~, a n = 30°, N A C A 6 4 A

r ! I 1 I

50 6 0 70 80 5 0 6 0 70 80 ( e ) Flap D.

(d) Flap F.

Figure 67.- The effect of the jet-momentum coefficient, cP, on the varia tion of the increment of lift coefficient with flap deflection; NACA 0006 airfoil section; s / c = 0.00110; tin = 3 5 O .

0 I O 2 0

30 40 5 0 6 0

70 80

Figure 68.- The effect of the jet-momentum coefficient, cP, on the variation of the increment of lift coefficient with flap deflection for the flap of reference 5; 0009-E4 airfoil section; slat position (10); s/c = 0.0050.

--

Theory

Ref. 9

0 Ref.16 Double slotted flap,

Cf =0.306~, 3 0 ' & , =

6 4 A O I O

0.1 2

.08

.06

.04 .02

30 40 5 0 60 70

( b ) S/C = 0.00036. (a) s/c = 0.00072.

Figure 69.- The effect of the jet-momentum coefficient, cp, on the variation of the increment of lift coefficient with flap deflection for the flap of reference 9.

NACA 64A010 airfoil section; flap position D; 6 , = 2 0 ' .

P w 0 --Theory Ref. 4

A Ref.17 Single slotted

f l a p , cf =0.257 c ,

2301 2

N A C A I I

I I

~'

-- I I-1-d-i

0 2 0 30 40 50 2 0 30 48 5 0

( b ) Flap type e; s l a t p o s i t i o n 6e; ( a ) Flap type f ; slat p o s i t i o n 6e; s / c = 0.0050. s / c = 0.00667.

Figure 70.- The e f f e c t of t h e jet-momentum c o e f f i c i e n t , cp, on t h e v a r i a t i o n of the increment of l i f t c o e f f i c i e n t with f l a p d e f l e c t i o n for t h e f l a p s of reference 4; NACA 23012-64 a i r f o i l s e c t i o n , b

---

Theory Ref. 1 2 cP

Single slotted - - - - I ) - - ~

5 A Ref. 17

0.28 !

flap, Cf N A C A . 2 0 . I 6 .I2

. O 8

.O 4

0 I O 2 0 30 40 50 60 70 80

( a ) s / c = 0.0090 Figure 71.- The e f f e c t of the jet-momentum c o e f f i c i e n t , cp, on t h e v a r i a t i o n of t h e increment P of l i f t c o e f f i c i e n t with f l a p d e f l e c t i o n f o r t h e f l a p of reference 12; NACA 23015 a i r f o i l w P section; no leading-edge h i g h - l i f t device.

---Theory Ref. I 2 I (b) S/C = 0.0070 Figure 71.- Continued.

I ---Theory Ref. 12 CU

------- A Ref.17 Single slotted

flap, cf =0.257c, NACA 23012 - - - , - , - . - _ _ I

-----

-----

- 3

a"

I

to 2 0

3 0

40 50

70 80

( c ) S / C = 0.0050 P Figure 71.- Continued.

w w

---

T h e o r y

GP

Ref. 12 I I I I I I I I

0.24 .I6 .I2

.-

A N

06 4

.04 0 2 I

30 40 5 0 60 70 80

IO 2 0

(d) S / C = O.OOl5 Figure 71.- Continued.

r

---

---- Theory

Ref. 1 2 A Ref. 1 7 Single slotted c f = 0 . 2 5 7 ~ , N A C A 23012 c P

4 0.12

.

. 0 8 e k .06 .04

.o 2

I

60 70 I 0

2 0 30 40 50

I O ( e ) s / c = O.OOO~ Figure 71.- Concluded.

I Except as noted,

6 c

Cf -

0.009 e-0.00110 and FlapA Ref. Airfoil Section 7 C .24

H i s against the nozzle

NACA 23012-64 .24 0.00667 , NACA 23012-64.25 0.005 .20

-----

5 0009-E4 .25 0.005 I

0 9 NACA 648010 .28 0*0f$36 7

0.00072 ; .I 6

--- 12 NACA 23015 .25 OS noted i,arrl O*Oo5

0.0015 d 0.00065 to 0.00017 - ; 0 ’ 40 50 60 70 0 IO 20 30 40 5 0 60 70 80 (a 1 P re s e nt investigation . (b) Reference investigations.

Figure 72.- The critical jet-momentum coefficients for the models of the present and the referenced investigations; a = 0.

Except o s noted,

Ref. Airfoil Section 9 ' A T h ' ' '

7 ~ = O . O O I I O and F l a p A - C

C o r y f o r Ref.

is against the nozzle 4 Flap NACA 23012-64 .24 0.00667 (type, e)

-___ 4 Flap NACA 23012-64 .25 0.005

6 TrFF; (type, f)

---- -

5 0009-E4 '25 0.005 0 9 NACA 64AOIO .28 0.00036

I / I 1 A

I- Theory far 40 50 60 70 0 IO 20 30 40 50 60 70 80 (a) Present investigation.

(b) Reference investigations.

Figure 73.- The theoretical increments of lift coefficient, and the measured increments at the critical jet-momentum coefficient for the models of the present and the referenced investigations .

Except a s noted, s

Ref. Airfoil Section 3 -

~ = 0 . 0 0 1 1 0 and Flap A C C is against the nozzle

H

----- 5 0 0 0 9 - E 4 . Z S 0.005

9 NACA 6 4 A O I O .28 0.0tn$36

0.00072 “ 1

---

I2 N A C A 23015 .25 os noted

I I I I I

40 50 60 70

a

a

(0) Present investigation.

(b) Reference investigations.

Figure 74.- The rate of change of the increment of lift coefficient with jet-momentum coefficient for values of the momentum coefficient greater than the critical value for the models of the present and the referenced investigations.

Except as noted,

Ref. Airfoil Section 2

--- ~=0.00110 and Flap A NACA 23012-64 z 0 . 0 0 6 6 7 is against the nozzle

----

- - 1 - l -

.36 -j-k----

-----

5 0009-E4 .25 0.005 ' 0 9 NACA 64AOIO .28 0 . 0 0 0 % and

--- 1 2 NACA 23015 .25 asnoted, -0,009

-32 --I---

- - I + - + -

.24 .20 L P 3.

.I6 . I 2 .08 .04 30 40 50 60 70 0 IO 20 30 40 50 60 70 80 8 8 (a) Present investigation.

(b) Reference inVeStigatiOnS.

Figure 75.- The jet-momentum coefficients required to achieve the theoretical increment of lift coefficient f o r the models of the present and the referenced investigations.

Figure 76. - Blowing-parameter r e l a t i o n s h i p s for s u b c r i t i c a l pressure r a t i o s .

.c Figure 77.- Blowing-parameter relationships f o r pressure ratios up to 1 0 .

I E c C ; E I c Q) Ptj

Pressure ratio, -

PO Figure 78.- The v a r i a t i o n with t h e p r e s s u r e r a t i o of t h e r a t i o of t h e momentum c o e f f i c i e n t for a convergent nozzle t o t h a t f o r a convergent-divergent nozzle.

9.0 8.0 7 . 0 6.0 5.0 4.0 3 . 0 2 . 0 I .o I .I 1.2 1.3 I .4 1.5 I 6 I a I .8 I . 9 hP

-

- P'I

X S W pto Figure 79.- Relationships among the blowing and power parameters for pressure ratios less than the critical value.

M=.117 airfoil section

a

0 Flop A. cx1md.d 0.00110 NACA 0 0 0 6 0 Ref. 12 0.0090 NACA 23015 0 Ref.12 0.0 00s NACA 23015 A Flop A agoinst nozzle 0.00110 NACA 0 0 0 6 V Flap A agoinit nozzle 0006 0.00065 NACA I 50 60 70 (0) Horsepower .04 .03 ca .02 .o I O 4 40 50 60 70 1 50 60 70 40 50 60 70 (b) Moss-f low coefficient 4 . 0 3 . 0 2 . 0 I .o 40 50 60 70 40 50 60 70 40 50 60 70 (c) Pressure rotio Figure 81.- Comparisons of the power, flow coefficient, and pressure ratio for the critical momentum coefficient for several blowing- flap arrangements of the present investigation and of reference 12.

NATIONAL AERONAUTICS A N D S P A C E ADMINISTRATION WASHINGTON. D.C. 20546 P O S T A G E A N D F E E S P A I D __- N A T I O N A L A E R O N A U T I C S A N D O F F l C I A L BUS IN E S S SPACE A D M I N I S T R A T I O N PENALTY FOR PRIVATE USE $ 3 0 0 SPECIAL FOURTH-CLASS RATE 451 U S M A I L BOOK 6 6 7 0 0 1 C 1 U A 7 6 0 7 2 3 S00903DS DEPT OF T H E A I R F O R C E AF WEAPONS L A B O R A T O R Y A T T N : T S H N I C A L L I B R A R Y (SUL) K I R T L A V D A F B NM 871 17 ..j ...

POSTMASTER If Undeliverable (Section 158 Postal hlnnual) Do Not Return ‘“The aeronautical and space activities of the United States shall be

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