WALL EFFECTS AND SCALZ EFF'ECTS W V/STOL MODEL TESTING

By the Staff of the

Powered-Lift Aerodynamics Section NASA Langley Research Center Langley Station, Hampton, Va.

Presented by

Kalman J. Grunwald*

This paper reviews the present status of knowledge with regard t o the limits of '

applicability of wind-tunnel data on V/STOL configurations, primarily with regard t o w a l l effects and scale effects.

The effects of the tunnel walls on the transition aerodynamics of a tilt-wing,

.

a fan-in-wing and a fan-in-fuse,lage configuration, as determined from t e s t s of a model of each configuration i n three different size tunnels are presented and discussed. The effects of the walls on pitcxing moment as well as l i f t and drag are included. reviewed.

Also, the applicability of Heyson's wall-effect theory 58

Comparison of small model results with large model or full-scale f l ight- test resul ts made where possible t o show where small-scale model data can o r cannot be used t o determine full- scale characteristics. figuration, XV5-A fan-in-wing configuration, and the Ames fan-in-fuselage config- uration are used. such comparisons is discussed.

Data fromthe VZ-2 con-

The overlap of wall effects with scale effects i n making

A brief discussion of the problems of ground simulation is included.

Nomenclature

model momentum area (for tilt wing based on sr(wing span)2

Efl(fan aameter)2 4

\

, for buried fan 4

r a t i o of Uf t ing systems (model) momentum. area t o tunnel cross-sectional area

model wing chord, f t

l i f t coefficient , L / ~ S

pitching-moment coefficients, My/@

thrust coefficient , T / ~ S

slipstream thrust coefficient , T/gsS

longitudinal force coefficient, X / ~ S

momentum coefficient, mVj/qS

drag, lb

l i f t , lb

pitching moment, f t - l b

number of fans

dynamic pressure, Ib/sq ft

supstream aynamic pressure, lb/sq ft

area, sq f t

fan thrust, l b

s t a t i c thrust, lb

velocity, ft/sec

I jet-exit velocity, ft/sec

ra t io of forward velocity t o hovering induced velocity \(Ref. 1)

velocity ra t io of forward velocity t o jet-exit

longitudinal force, lb

angle of attack, aeg

f lap deflectioqangle, deg

louver deflection, deg

/change i n downwash angle, deg

change i n angle of attack, deg

tip-speed rat io, V/UB

rotational speed,\ radians/sec

Subscripts:

C corrected * Aerospace Techiiologist , - Powered-Lift

Aerodynmics Section, Full-Scale Research Division. fuselage

-8-

h = a out of ground effect

W wing

t t a i l

S-ry

The results of recent investigations t o deter- mine the w a l l effects on tilt-wing and buried-fan configurations indicate that significant w a l l effects can be encountered with both types. Heyson's w a l l correction theory was found t o be adequate for correcting the l i f t and drag data of the two buried configurations investigated. tilt-wing configuration , Heyson's wall" correction theory was adequate i n correcting l i f t and drag t o the free-air case on ly i n the intermediate angle- of-attack ranges. A t the low angles of attack the corrections were ineffectual i n compensating for the wall effect, while a t very high angles of attack and very high drag conditions the drag correction was overestimated.

For the

The w a l l effects on the tail-on pitching moments for fan-in-wing configurations were found t o be of opposite sign t o that predicted by theory, thus indicating a need for a basic revision t o the theory as regards the wall-induced upmsh downstreas of the l i f t i ng element.

Scale effects on propeller-driven configura- tions such as the Vertol VZ-2 tilt-wing airplane have generally been found to be negligible. Com- parison of data from buried-fan configvrations, however, shows significant but inconsistent differ- ences between model and full-scale data. The results imply a significant effect of Reynolds num- ber. It i s also evident that extreme care must be exercised i n the simulation of the fan and that accurate measurement of the fan in l e t and exit momentum i s necessary.

Introduction

Wind-tunnel test ing of models has always been subject t o some limitations due t o the a r t i f i c i a l boundary conditions created by the tunnel walls and due to Reynolds number effects. airplane configurations these limitations have long been recognized. mize the scale and wall effects or the data are ut i l izea with ?hll recognition of the limitations due to scale and wall effects.

With conventional

Model tes t s are planned t o mini-

The old experience i s not always sufficient for V/STOL testing, however, for several reasons. First , the w a l l effects are altered by the large downwash angles associated with V/STOL configura- tions, and, secondly, there are new and unknown scale effects involved with some of the V/STOL configurations.

For unpowered conventional airplane configura- tions the classical wall correction theory, which assumes zero deflection of the wake from the m o d e l , i s adequate because the actual wake deflection i s very small. however, slipstream or jet deflections of the order of 900 can be encountered and these large wake deflections can produce large w a l l effects unless the model i s appreciably smaller than would be used i n unpowered model tes t s .

With powered V/STOL configurations,

Heyson (Ref. 1) has

developed a w a l l correction theory i n which these large w a h deflestkms aa?e %ken into aeemmt.

The purpose of this paper i s t o present sane results of recent MASA investigations t o determine experimentally the wall effects encountered i n several Mfferent size t e s t sections by a tilt-wing and two buried-fan configurations and t o check the applicability of Heyson's tunnel wall correction theory.

Some recent information on scale effects i s also included and i s presented a s comparison of moael and full-scale data on the Vertol BZ-2 tilt- wing airplane configuration and two buried-fan research configurations.

Wall Effects

Tilt-Wing Model

effects on tilt-wing configurations was tested i n three t e s t sections (the Langley 3- by &-foot full-scale tunnel, and the 7- by 10-foot, and 12.7- by 17-foot t e s t sections of the Langley J004PH 7- by 10-foot tunnel). instal led i n the IT-foot t e s t section i n Fig. 1. The wing model had a 4-foot span, an 18-inch chord, a d used an NACA 0015 a i r fo i l . A 30-percent-chord full-span Fowler flap was used on the model for the t e s t s discussed i n this paper. Each of the two propellers were 2 feet i n diameter. The model had no fuselage, but a &-inch section a t the center of the wing, from maximum thickness a f t , was contoured t o allow for the entrance of the sting and balance. The same balance, support sting, and readout equip- ment were used i n each of the three t e s t sections.

The model used i n the investigation of w a l l

This model i s shown

Experimental data.- A sample of the basic data

In spite of an attempt t o set the same without w a l l corrections applied i s presented i n Fig. 2. propeller thrust i n each tunnel, some variation was experienced, as shown, which necessitated cross- plotting the data t o obtain a comparison a t constant thrust conditions (Fig. 3 ) . Because of the very large size of the 30- by 6O-foot t e s t section with respect t o the model, the data from the 30- by €&foot test section are essentially free of w a l l effects and therefore assumed t o represent the free- air condition.

The data of Fig. 3 indicate that the wall effects for this model are small, but the curves do not f a U i n the expected order. The data indicate a wall-induced loss i n l i f t for the model i n the 7- by 10-foot t e s t section as expected, but the data from the 13.7- by 17-foot t e s t section indicate a higher l i f t which was not expected. for t h i s increase i n l i f t has been found although there has been speculation that it may be associated with the short length or unusual configuration of the t e s t section which i s followed by a contracting rather than an expanding section.

No explanstion

(See Fig. 1.)

Application of wall corrections.- The tunnel- w a l l correction theory developed by Heyson (Ref. 1) predicts both vert ical and horizontal induced velocities due t o the tunnelwalls. mese combine t o produce changes t o both angle of.attack and the free-stream velocity or dynamic pressure. of the predicted angle of attack and dynamic pres- sure corrections applicable t o the 7- by 10-foot test-section data of Fig. 3 are presented i n Fig. 4.

Examples

-9-

Application of these correction factors t o the data as shown i n Fig. 5 resul ts i n what appear t o be large corrections t o l i f t and drag. However, t he thrust coefficient must also be corrected as shown i n Fig. 5, and, as a resul t , there i s no longer a direct relationship between the corrected and uncor- rected data. I n order t o obtain a comparison of corrected and uncorrected data a t constant thrust coefficient, it i s necessary t o cross-plot the data. When such a cross plot i s made it i s found tha t the various corrections combine i n such a manner tha t the net effect i s a relat ively small wall correction as shown by the comparison of corrected anduncor- rected data presented a t constant thrust coefficient i n Fig. 6.

The data of Fig. 6 indicate tha t i n the in te r- mediate angle range somewhat be t te r agreement with the free-air condition (30- by 60-foot tunnel t e s t s ) i s obtained by applying Heyson's wall corrections t o the data from the 7- by 10-foot tunnel. However, a t the low angles of a t tack the corrections were ineffectual i n compensating for the wall effect , while a t very high angle of attack and very high drag conditions the drag correction was overesti- mated. diverge occurs a t a condition producing a w a l l - induced angle-of-attack correction of about 8 O at a thrust coefficient of 6 and about 180 a t a thrust coefficient of 14. Thus there does not appear a t present t o be any easi ly defined rule of thumb fo r the l i m i t of application of Heyson's theory corre- sponding t o the 2 O l i m i t suggested i n Ref. 2 i n connection with classical theory.

Buri ed-Fan Configurations

configuration shown i n Fig, 7 w a s a l/g-scale model version of the full- scale model tes ted i n the Ames 40- by &-foot tunnel (Ref. 3). The wall-effects investigation w a s conducted i n two t e s t sections; the 15.7- by 17-foot t e s t section and a 4.4- by Y-foot t e s t section ins ta l led in the 3OO-MPH 7- by 10-foot tunnel. w a s intended t o approximate a l/g-scale model of the 40- by 80-foot t e s t section.

The point a t which the drag data begins t o

Fan-in-fuselage model.- The fan-in-fuselage

The 4.4- by "-foot t e s t section

In general, the results of the wall-effect investigation showed a significant increase i n lift (Fig. 8) i n the smaller test section but negligible w a l l effects on drag and pitching moment ( for this tai l- off model). l7-foot t e s t section were estimated and found t o 'be negligible. increme i n l i f t is due t o a wall-induced upwash (of about 20 to 4O i n Fig. 8(a)) which i s reasonably w e l l predicted by Heyson's theory. t o free-strea aynamic pressure were found t o be much smaller tbaJl f o r the tilt-wing model asd because of the m e r of presenting the data .these corrections cixmge on= the velocgty r a t i o v/Vj.

The corrections i n the 15.7- by

As can be seen from Fig. 8(a), the

The corrections

Fw-in-wlraa model; l i f t and drag.- The f a - i n - wing =del {F%& 9 ) $sed i n the wall&fect an& scale-effect investigation was a 0.18-ecale model of the Geaeral Electr ic - Ryan XQA airplane. ThLs model wes tes ted in three tes t sections; the 30- by @-fQot t e s t section (which can be considered tQ represent free-air CondiWons) , and t.hs 7- by lO&oot ami L5*7- by lT-Taot test sections. The resul-ks obtained (Fig. le) are similar t o the wall effects on the fan-in-fuselage model i n tht sub- s'cantially higher lifts were measured i n the smaller

t e s t sections and the w a l l effects on drag were negligible. Application of Keyson's corrections brought the data from the three t e s t sections into reasonably good agreement.

Pitching moments.- O f the three models used i n wall-effects investigations only the fan-in-wing model had a horizontal t a i l and therefore could be used t o obtain data relat ive t o the wall effects on the flow f i e l d i n the region of the t a i l (Fig. 9 ) . Data from t h i s model (Fig. 11) i n the 7- by 10-foot t e s t section indicate tha t the tunnel walls pro- duced a nose-down increment of pitching moment indicating a wall-induced reduction i n the downwash angle a t the t a i l . Application of w a l l corrections from Ref. 1, however, does not correct for t h i s wall effect but moves the curves further apart. mis i s because Ref. 1 predicts a smaller wall- induced upwash angle a t the t a i l than a t the wing.

The apparent reason fo r the inadequacy of the theory i s i l lus t ra ted i n Fig. 12. The theory assumes tha t the wake from the l i f t i n g element (which i n t h i s case i s the slipstream from the fan) continues i n a s t raight l i ne t o the tunnel floor. I n practice, however, t h i s wake i s deflected down- stream by the free-stream flow (Ref. 4). It i s interesting t o note tha t the calculated variation of wall-induced upwash reaches a maximum a t the point where the assumed wake impinges on the floor and decreases downstream of t h i s point. In classi- c a l w a l l correction theory, on the other hand, the wall-induced upwash continues t o increase down- stream of the l i f t i n g element and a t in f in i ty reaches a value double tha t a t the l i f t i n g element. The actual variation of wall-induced upwash with a curved wake i s not known but it may follow a varia- t ion similar t o that shown i n Fig. 12. Additional work i s needed t o develop a w a l l correction theory tha t w i l l adequately predict the effects of the tunnel walls on the flow f id ld i n the region of the t a i l .

Choice of Model Size

V/SML model i s a function of several variables such as the precision of the data required, the avai labi l i ty of equipment such as model motors or compressed a i r , and the cost and t i m e required for model construction and tunnel testing. are involved only i n the f i r s t of these.

The decision as t o the appropriate size for a

Wall effects

An approximate indication of the model size needed t o avoid appreciable w a l l effects Tor buried- fan and jet-v/sWL configurations i s presented i n Fig. 13. The parameter tha t determines the magni- tude of the wall effects i s the r a t i o of tbe momen- 'cum area o f %he model &, t o the tunnel cross- section area AT. The momentum area of the model I s the cross-sectional area of the wake from the l i f t i n g element and i n %he case of j e t and buried- f a confi,gura%ions at low speeds where most of the Ufi i s c&rr2ed on the fans o r jets t h i s 8rea should. be taken as the fan or je t area. derived f o r the case of ver t ica l Je ts , are turned af t the w a l l effects are reduced, that i s , the bounaaries shown i n Fig. 15 move upwara. Nevertheless, Fig. 15 indiczbes tha t if apprecLable w a l l effects are t o be avaided wit& je% and buried- fan C O ~ f ~ g U r a ~ ~ Q ~ S %he jet o r fan area must be kept cwsidez-abv below 1 percent o f fhe tunnel cross- sectional area.

Fig. 13 i s As the j e t s

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Fig. 13 does not apply t o ti l t-wing and deflected slipstream configurations. With these configurations the l i f t i s usually spread more evenly over the wing fo r most of the range of t e s t conditions of interest and therefore the appropriate momentum area i s tha t of the c i rc le circumscribing

the wing span The momentum area r a t i o

fo r the tilt-wing model i n the 7- by 10-foot t e s t

section was = 0.18 . It i s f e l t that i n general

a momentum area of 10 percent or l e s s of the tunnel cross-sectional area would be preferred for most tilt-wing and deflected slipstream configuration investigations

Am = - . ( f) (4 )

Scale Effects

Propeller Configurations

been bui l t there has been relat ively l i t t l e corre- sponding model data available t o make direct com- parfLsons with f l ight- test resul ts . t e s t beds were bu i l t with l i t t l e or no wind-tunnel

Although several V/STOL t e s t bed a i rc ra f t have

Most of t he

support.

A n example of some of the limited data tha t are available i s presented i n Fig. 14. The good agreement between the quarter-scale model (Ref. 5 ) and f l ight- test (Ref. 6) resul ts indicates tha t there i s l i t t l e effect of Reynolds number on the l i f t , drag, and pitching moments i n steady leve l f l i gh t transition. Similar indications of negligi- ble scale effects have also been found by the French on the Breguet.gk0 and 941 deflected slipstream a i r - craf t developments.

One of the primary problems involved i n tilt- wing configurations i s that of predicting the descent and deceleration capability of the airplane as discussed i n Ref. 7. There is, however, no ade- quate means currently available t o predict, from l i f t-drag polars, the limiting amount of s t a l l that the p i l o t w i l l to lerate i n descending and deceler- ating f l igh t and therefore no adequate means of determining the maximum deceleration or descent angle capabilit ies of the airplane.

Euried-Fan Configurations

buried-fan configurations used i n the wall-effects investigations are available from t e s t s i n the Aaes 40- by &-foot tunnel (Refs. 3 and 8). These data (corrected for wall effects) are compared with the model data (also corrected f o r w a l l effects) i n Fig. 15 and indicate tha t there can be significant scale effects on buried-fan models and tha t not only the magnitude but the direction of the apparent scale effect i s a function of the configuration, A s shown i n Fig. 15, the l i f t of the fan-in-fuselage model i s smaller than f u l l scale while for the fan- in-wing configuration the model l i f t i s greater.

Full-scale data on models corresponding t o the

There are several parts of a buried-fan con- figuratian, as shown i n Fig. 16, that may be sub- jec t t o Reynolds Dumber effects. (I-) the Wing leading edge which i s always sensitive t o Reynolds number effects and i s particularly so i n a buried-fan configuration because of the large upwash induced a t the leading edge by the fan; (2) the fan i n l e t which nay separate i f too small

These are:

a radius i s used and i s probably sensitive t o Reynolds number; (3) -&e fm lltself which, experi- ence with small-scale propellers indicates, should exhibit somewhat altered characteristics a t l o w Reynolds numbers; (4 ) the exi t louvers which may exhibit higher losses a t low Reynolds number; ( 5 ) the pressures induced on the lower surface of the model by the exiting stream which may be sub- ject t o Reynolds number effects.

I n general it appears tha t a bet ter under- standing of the basic aero-ics of buried-fan configurations i s required before the effects of Reynolds number can be sorted out.

An important factor i n determining the aero- dynamics and scale effects of buried-fan config- urations i s the adequate simulation of the i n l e t and exi t momentum. The model of the fan-in- fuselage configuration did not use a scaled fan but had the fan, i t s housing, and the i n l e t out t o the l i ne shown i n Fig. 7 mounted on s train- gage beams i n an attempt t o measure the fan thrust. Fig. 1-5 are nondimensionalized by dividing by t h i s thrust. The full- scale thrust was determined by integration of surveys of the exi t flow. A t l eas t part of the differences shown could be due t o differences i n the method of determining the fan thrust.

The lift and drag data presented i n

No attempt was made t o measure the thrust of the fans i n the fan-in-wing model but the fans were bu i l t t o scale as accurately as possible, including the use of a t i p turbine drive which was powered by compressed a i r . Inasmuch as the thrust was not measured, both the model and full- scale data are nondimensionalized by dividing by the s t a t i c thrust and plotted against the tip-speed rat io . Although the model fan was bu i l t t o scale there i s no assur- ance tha t the variation of fan thrust with forward speed i s the same as fo r the full- scale fan. A s indicated above, scale effects on fan character- i s t i c s are t o be expected.

Ground-Effect Simulation

A number of studies have indicated tha t STOL configurations operating a t high l i f t coefficients (deflected slipstream as wen as j e t f l ap) can suffer serious losses of l i f t i n ground effect. The investigation reported i n Ref. 9, however, indicates tha t the losses measured by a model over a conventional fixed groundboard i n a wind tunnel may be appreciably greater than would be experi- enced by an airplane moving over the ground (Fig. 17). The moving model resul ts were obtained by mounting the model on a towing tank carriage t o move it over the groundboard.

The reasons fo r the differences i n resul ts have been graphically i l lus t ra ted by some recent French flow visualization studies using the water tunnel (Ref. lo). In these studies (Fig. 18) the moving model w a s simulated by using an endless be l t t o siaulate the ground and moving the be l t at t he same speed as the water. studies (Fig. 18) showed tha t with the moving ground the j e t sheet from the jet f lap did nat dis- turb the flow very far upstream of the point of impingement; whereas with the fixed ground the boundary layer between the ground and the free- stream flow w a s separated well forward of the model thus al ter ing the ent i re flow f ie ld .

The flow visualization

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Because of the importance of ground effects t o V/STOL aircraft a moving-belt groundboard system (Fig. 19) i s being obtained for the l7-foot t e s t section of the Langley JOO-MPH 7- by 10-foot tunnel.

Concluding Remarks

The wall-induced l i f t and drag errors experi- enced by the tilt-wing model and two buried-fan models could, i n general, be adequately corrected by Heyson's method. The tail-on pitching-moment data, however, indicated a higher wall-induced upwash a t the t a i l than a t the wing, whereas the theory indicated a lower wall-induced upwash a t the t a i l than a t the wing thus giving the wrong sign t o the pitching-moment corrections. Additional work i s needed t o develop an adequate theory t o correct the pitching-moment corrections.

Scale effects on the propeller-driven VZ-2 configuration i n which the wing i s largely immersed i n the propeller slipstream have been found t o be small. However scale effects on the buried-fan configurations are of significance. crepancies caused by scale effects on the buried- fan configurations appear t o be of greater magnitude than those caused by w a l l effects. The need t o develop means t o adequately simulate the fans and accurately measure the in l e t and exit momentum is of prime importance.

The data dis-

TEST SECTION

a, DEG

References

1. Heyson, Harry H., "Linearized theory of wind- tunnel jet-boundary corrections and ground effect for VTOL-STOL aircraft," NASA TR R-124 (1962).

2. hscombe, A . , and Wi l l i ams , J., "Some comments on high- lift test ing i n wind tunnels with particular reference t o jet-blowing models," Rep. 63, AGARD North Atlantic Treaty Organiza- t ion (Brussels) (Aug. 1956).

3 . Aoyagi, Kiyoshi, Hickey, David H., and desavigny, Richard A., "Aerodynamic character- i s t i c s of a large-scale model with a high disk-loading l i f t i n g fan mounted i n the fuse-

- lage," NASA !l?N D-775 (1961).

4. Jordinson, R., "Flow i n a j e t directed normal t o the wind,," R ana M No. 3074, ARC TecMcal Report. Ministry of Supply, Aero. Research Council.

5. Newsom, W i l l i a m A . , Zr . , an& Tosti, Louis P., "Force-test investigation of the s tab i l i ty and control characteristics of a l/k-scale model of a tilt-wing vertical-take-off -and- landing aircraft ," NASA Memo =-3-58~.

6. Pegg, Robert J., "Summary of f l igbt- test results of the VZ-2 tilt-wing aircraft," NASA mJ D-989 (1962) *

7. Mitchell, Robert G., "F%iLl-scale wind-tunnel t e s t s of the VZ-2 VML airplane with particular reference t o the wing s t a l l phenomena," NASA m D-2013 (1963).

8. Kirk, Jerry V., Hickey, David H., and Hall, Leo P., "Aerodynamic characteristics of a full-scale fan-in-wing model including results i n ground effect with nose-fan pitch control," NASA Proposed TN.

9. Turner, Thomas R., "Ground influence on a model a i r f o i l with a jet-augmented f lap as determined by two techniques," NASA TN D-658 (1961).

10. Werle, Henri, "Simulation D i L'Effet De Sol Tunnel Hydrodynamique. Office National D'E-tudes Et De Recherche6 'Aerospatiales," T.P. n- 63 (1963).

Au

4 A55 I (=FREE AIR)

I L7' X IO'

0 2040M)8010 8 6 4 2 0 - 2 - 4 - 6 - 8 a, DEG cx

-12-

"r

4 I I I I I I I I

.7L , I I I I I I I I

0 1 0 x ) 3 0 4 0 5 0 6 0 x ) 8 0 9 0 CORRECTED ANGLE OF ATTACK, a + Aa, DEG

I l l ]

2o r I rAa=18"

\301 X 60' (-FREE .AIR)

x IO' UNCORRECTED

21- I L L

0 20 40 60 80 IO 8 6 4 2 0 -2 -4 -6 -8 -10 cx OR CX, a OR ac

Figure 6.- Corrected and uncorrected data. coqared at constant t h s t coefficient with free air condition.

1 I I I I I

1.5

L r,

TEST SECTION 0 15.7'x17'sjFREE AIR 0 4.4'X7' UNCORRECTED d 4.4'X7' CORRECTED

BY REF. I .5L -4b -5 ; Io 1; 20

ANGLE OF ATTACK, a, DEG

(a) Effect of an& of attack, B 0.32. v3

Figure 8.- wall effects on fan-in-fuselage madel.

35.36- FT F U L L SCALE

!--& 5 . 2 5- F T F U L L S C A L E 7.0- IN. MODEL

1.5 15.7'~ 17' TEST SECTION 4.4'x7' TEST -

SECTION " LI UNCORRECTED

IJ FREE AIR 7 -- -- __ - '"" '- L

TF <CORRECTED I BY REF. I

'5 L I I I I 1 I

0 . I .2 .3 .4 - 5 .6 V/Vj

(h) Effect of velocity, a = Oo.

Figure 8.- Cmcluded.

-13-

.I8 SCALE

I .Qk

Figure 9.- Fan-in-wing model used in wall-effects investigation.

D - TS

--* 7'X IO' CORRECTED --a 15.7'~ 17'1 BY REF. I 0

2.0 r I

(a) Effect of angle of attack, p 0.22, = 0.48. vj,static

Figure 10.- Wax1 effects on lift and drag of fan-in-wing model.

0 D5 .IO .I5 -20 ,25 TIP SPEED, RATIO, p

(b) Eifect of velocity, a - 00.

F i g v r e 30.- concluded.

V- -x-i-JY

1 .-. '\PROBABLE VARIATION

- AM AT

30' X 60' WFREE AIR

7 X 10 UNCORRECTED

I -40 -4 A b Ib I; 20 85 ANGLE OF ATTACK, a, DEG

FAN-IN-WING MODEL IN - 7 ' X IO'

.02

FAN-IN-FUSELAGE MODEL .015

LIFT ERROR DUE TO WALL EFFECT

,010

-5% ,005

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

w = 3500 L!3 I00

WING i,+arf, OFATTACK, ANGLE DEG 6o I\ 40

20

0 FLIGHT TEST REF 6 ’

0 SCALE MODEL REF 5

I I I I I I 0 , 0 20 40 60 80 I00 120

TRIM LEVEL-FLIGHT SPEED, V, KNOTS ( a ) Hing angle of attack.

F i w e 14.- CmBpImn of transition chamcteristics of Vert01 Vz-2 tllt-uing airplane as obtained f*Drnmodel and full-scale testa.

a = 0; LOUVER = 0; CORRECTED FOR WALL EFFECTS

FAN-IN-FUSELAGE CONFIGURATION

FAN-IN-WI NG CONFl GURATI ON

FULL SCALE

TF 1.0

.4 .6 V - v j

1.5 MODEL :-”

TS

. 5 ~ .I .2 .3 P

l.mGmNAL

PERCENT OF TOTAL TRAVEL

STlCK POSITION, -0

(b) Btick posit ion rcqvired for l c m g i ~ n a l trim

Figure 14.- Concluded.

LOWER -SURFACE EX’T SUCTION PRESSURES

Figure 16.- Scale effect problem areas on buried-fan configurations.

Figvre 15.- Comparison of scale effects on fan-in-fuselage and fan-in-wingmodels.

REF. 9 LIFT IN GROUND EFFECT

LIFT OUT OF GROUND EFFECT

1.0 I ---- *,,/,? ------ HNIQ!E

.%

/ TUNNEL TECHNIQUE .6 CONVENTIONAL WIND

0 I 3 4

Figure 9.- Effect of t e s t technique on the loss i n l i f t experiencedby a jet- flap model i n pound effect.

-15-

REF. IO

B 0 ' - 0 0 0

MOVING BELT GROUND BOARD

FIXED GROUND BOARD

Figure 18.- Effect of ground simulation technique on the flow f i e ld aroma a two-dimensional je t- f lap model.

MODEL PIVOT POINT AND CENTER OF INTERNAL STRAIN GAGE BALANCE

Figure 19.- Proposed ins ta l l a t ion of the moving b e l t groundboard i n the 17-foot t e s t section of the NASA Langley 3OO-MPH 7- by 10-foot tunnel.

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