Large-scale wind-tunnel tests of descent performance of an airplane model with a tilt wing and differential propeller thrust

Document

GASE FILE

CORY

NASA TECHNICAL NOTE NASA TN D-1857

GO '-7, Z I--.

LARGE-SCALE WIND-TUNNEL TESTS

AN

OF DESCENT PERFORMANCE OF

WING

AIRPLANE MODEL WITH A TILT

PROPELLER THRUST

AND DIFFERENTIAL

Robert Page, by Wallace H. Deckert, V.

and Stanley O. Dickinson

Ames Research Center

Moffett Field, Calif.

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION • WASHINGTON, D. C. ° OCTOBER 1964 LARGE-SCALE WIND-TUNNEL TESTS OF DESCENT PERFORMANCE OF AN AIRPLANE MODEL WITH A TILT WING AND DIFFERENTIAL PROPELLER THRUST By Wallace H. Deckert, V. Robert Page, and Stanley O. Dickinson Ames Research Center Moffett Field, Calif.

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sole by the Office of Technical Services, Department of Commerce, Washington, D.C. 20230-- Price $1.00 LARGE-SCALE W!ND-TUKNEL TESTS OF DESCENT PERFOPJ4ANCE OF A]_ AIRPIF_E MODEL WITH A TILT WING A_,YD DIFFERENTIAL PROPELLER THRUST By Wallace H. Deckert, V. Robert Page, and Stanley O. Dickinson AI_es Research Center Moffett Field, Calif.

SU_4ARY Tests were conducted to determine the wing stall, performance, and longitudinal stability and control characteristics of a large model of a V/STOL tilt-wing transport aircraft. The scope of the tests was limited primarily to the low-speed transitional regime.

Test configurations included wing tilt angles from 0 ° to 40 ° , double- O slotted trailing-edge flaps deflected from 0 ° to O0 , various wing leading- edge devices; such as partial-span and full-span Kr<[ger flaps and slats, and several rsmps that extended from the top of the fuselage to the tilted wing center section.

Test results show that air-flow separation over the tilted wing center section limited maximm_ lift and produced horizontal-tail buffet, that theoretical wing stall boundaries were optimistic, that low-speed descent angles were less than 5 ° for the various flap, slat, and rm_p configurations tested_ and that descent angles greater than i0 ° were obtained by operating the inboard propellers at a higher thrust level than the outboard propellers.

Pitch control with the variable incidence horizontal stabilizer was sufficient for tri_ed i g level flight throughout the transition range tested to a minimum airspeed of 40 knots. An additional pitch control device, such as a tail rotor, would be required for maneuvering and steady c!i_ing flight, including the landing wave-off condition for a configuration with no wing tilt.

The tail on and off longitudinal static stability was slightly unstable throughout the transitional flight regime tested.

INTRODUCTION NASA investigations of the aerodynsm_ic characteristics of propeller- driven tilt-wing aircraft have included wind-tunnel and flight tests. The results of the wind-tunnel tests of the first large-scale, four-propeller,

tilt-wing model (refs. i and 2) are of particular interest for a background

to this report. In references i and 2 it was i_dieated that air-flow separa-

tion on tilt-wing aircraft would limit descent performance and cause buffet-

ing in the low-speed transitional flight regime. Flight tests of the VZ-2

tilt-wing ai_lane (ref. 3) also indicated that th_ most critical region of

operation of the VZ-2 was during decelerating conversion and/or descent where

the effects of air-flow separation and wing stall Sed to buffeting, erratic

motions, and general deficiencies in handling qualities. For the model of

references i and 2, premature air-flow separation occurred from the tilted

wing center section outside the propeller sliDstream and from an outboard

area of the wing between the nacelles which was o_ly partially immersed in

the propeller slipstream.

The objectives of the wind-tunnel investigation of this report were (i)

to determine whether the problems revealed in references i and 2 were present

on a model which closely resembled a V/STOLtilt-wing transport aircraft, (2)

to determine the effects of various spanwise extents of leading-edge devices

on the onset of air-flow separation, and (3) to c]et_.rmine descent performance with the inboard propellers operating at high thrust relative to the outboard propellers.

NOTATION A total disk area of all four propellers, _ sq ft b wing span, ft c wing chord parallel to plane of s_nmetz_, _, _T't 2 ob/2 c mean aerodyns_nic chord, _ c 2 dy, ft _O measured dra_ CD drag coefficient including thrust_ Q_S measured lift lift coefficient including thrust, CL cl S slope of lift curve, per deg CLc_ itchin c moment pitching-moment coefficient, Cm $Cm %i t D propeller diameter_ ft acceleration of gravity, 32.2 ft/sec 2

g

incidence of horizontal tail relative to fuselage reference line, deg

it

incidence of wing root chord with respect to fuselage reference line with wing down, deg (0 for model of this report) v_ J propeller advance ratio, L lift including thrust component3 ib n propeller rotational velocity, rps

%

free-stream dynamic pressure 3 Ib/sq ft pv_ R Reynolds number, r propeller blade radius, ft S wing area, sq ft T total thrust of all four propellers, ib T T' C thrust coefficient, q S V airspeed scaled for W/S = 7 0 ib/ft 2 airplane, knots descent rate_ ft/min VS V free-stream tunnel velocity, fps or as noted oo W gross weight, ib U_ angle of attack of fuselage reference line, deg geometric angle of attack of wing root chord 3 _ + _ + iw, deg % propeller blade angle at 3/4 r, deg 7 descent angle_ deg flap deflection relative to local wing chord3 deg 5f 5 w wing tilt angle of root chord relative to fuselage reference line 3 deg coefficient of viscosity, slugs/ft-sec mass density of air, slugs/ft 3 P

Relationships of Coefficients Based on Free-Stream

and Propeller Slipstresml Dynamic Pressure

T

% :%+K _7 T T C _ c CTs qs A ' T c + A/ T c _ !. 38 i - CTs - qs CLs : CL(1 - CTs) MODEL AND APPARATUS The model installed in the center of the 40- by 60-foot test section is shown in figures i and 2. The model resemble_ v_2 V closely the XC-142 V/STOL airplane at 0.6 scale.

A three-view drawing of the model is sho_m in figure 3. Pertinent geo- metric characteristics (table I) of the wing w_::_'e: span, 40.5 feet; aspect ratio, 8.4; taper ratio, 0.5}; and mean aerodyn&mic chord, 4.99 feet. The wing airfoil section for the model (table II) was an NACA 23017 with the leading edge modified to the geometry of an NACA _]3-318 section. (The airfoil _<_ ._7<, section.) section of the XC-142 airplane is an NACA ..... .]-)_,,_ A typical section of the double-slotted trailing-edge flap is shown in figure 4; the coordinates are presented in taLlc ii. The large trailing-edge flap extended full span, except for nacelle c _touts, and was deflected a maximum of 60 ° • Typical sections of the various wing leading-edge devices that were evaluated are shovau in figure 5. The wing leading-edge configurations included partial-span and full-span extents of a O.lOc KrUger flap and 0.2c slats. The partial-span leading-edge devices extended from 43.7- to 60.7- percent and from $6.4- to 100-percent semispan.

The geometric characteristics of various r_os that extended from the top of the fuselage to the tilted wing center section are shown in figure 6.

The basic short fore and aft ramps were installed for most of the runs. The modified fore and aft ramps were about twice the length of the basic fore and aft ramps, respectively.

The geometric characteristics of the three-%laded model propellers are shown in figure 7. The solid aluminum model propellers had a diameter of 9.3 feet, an activity facto_ A.F., of 121 per blade, and rotation was such that the outboard blade of all four propellers rotated upward. Each propeller was Shaft mounted on a gear box and driven by an electric motor. The four motors were operated in parallel from a variable-frequency power supply.

TEST AND PROCEDURE oo knots Tests were made at free-stream tunnel velocities from 17 to F- (_o : i to 15, Reynolds number 0.9 to 3.6 million based on the wing mean aerodynamic chord of 4.99 ft). For each run the fuselage angle of attack was varied while the tunnel dynamic pressure, propeller speed, and propeller blade angle were held fixed.

The three-component longitudinal data presented in the figures include the propeller thrust components as well as the model aerodynamic components.

The propeller thrust characteristics were determined by wind-tunnel tests at a thrust axis inclination of 0 ° with the propeller on and off the model. Fig- ' for this ure $ shows the relationship of the thrust coefficients T c and CTs model so that the coefficients based on free-stream dynamic pressure may be readily converted to coefficients based on propeller slipstream dynamic pres- sure as desired.

Pitching moments are presented about the center sho}_ in figure 9. The moment center used for calculation was varied slightly with wing tilt.angle in accordance with a center-of-gravity change due to wing mass effects expected for a tilt-wing transport aircraft. For a 0 ° wing tilt angle, the horizontal moment center was at 0.23 _.

The model was mounted on faired struts in a manner which isolated the strut forces from the model. Tunnel wall corrections and strut-to-model interference factors were not applied to any of the data.

RESULTS Figures i0 to 26 are the general results of configurations having partial-span F_ger flaps, full-span slats, and the combination of full-span slats with differential spanwise propeller thrust.

DISCUSSION Aerodynamic Characteristics of the Basic Configuration _hree component longitudinal force and moment data, CL, CD, and Cm, are presented at several thrust coefficients in figures i0 and ii for the basic configuration for wing tilt angles, _w, of 20 ° and 30o and flaps deflected 60 ° . The basic configuration consisted of partial-span F_ger flaps_ the basic short fore and aft ramps over the wing cente_ _ section (fig. 6), a blade angle setting_ _ of i0 ° on all four propellers_ and a horizontal-tail setting of +20 ° • !

At T c > 0_ maximum lift was limited by air-flow separation from the portion of the tilted wing center section that _as unprotected by the propel- ler slipstresza and from an area on the wing between the inboard and outboard nacelles. As shown on figures i0 and ii_ the air-flow separation limited the descent capability of the basic configuration t_ descent angles of about 2 ° or 3 ° • This air-flow separation originated on the _pper surface of the _ing just aft of the leading edge and at a spanwise ]oc_tion above the sides of the fuselage. As the fuselage angle of attack _as increased, the separated area progressed aft to the wing root trailing edge and outboard into the area immersed in the inboard propeller slipstream and along the upper surface of the fuselage to the empennage. At the higher fuselage angles of attack, the stalled center section wake enveloped about $1Jo-thirds of the vertical tail.

The air-flow separation at the wing center _3ection was generally asso- ciated with buffeting of the horizontal tail. q_ical buffet intensities, measured at the tip of the horizontal tail_ a_'o _h_n in figure 12 for the basic configuration with a wing tilt angle of 20 °. Figure 12 shows that the buffeting level increased rapidly as the fuselag_ angle of attack was increased above 4 ° • _ms the test result (figs. IC to 12) supported the gen- eral conclusions of references i and 2 that th_ flifht envelope of a tilt- wing aircraft with an exposed wing center s_cti_n uay be limited by separation over the wing center section.

Incremental flap lift and drag coefficiei_t_ a_'._+ _ presented in figure 13 for the 0° untilted wing configuration. At con_a_t thrust_ the flap lift coefficient increased approximately linearly f_om D° to 60 ° flap deflection.

The flap drag coefficient remained unchanged f_'_,m yo to about 30 ° flap deflec- tion and increased rapidly from 30 ° to 60 ° derl_ction. Figure 13 indicates that both lift and drag would increase if the fiascos were deflected beyond 60 °.

Hence_ flap deflections greater than 60 ° may l_e _irab!e for higher descent angles during STOL t_q0e landings. Vane and flaw_ _ _+rface tuft observations and surface pressure data_ corresponding to test c _ditions of figure 13, indicated that the flow over the surface of all th = vanes and inboard flap was attached but the flow over the outboard f_aos was separated. The area of attached flow over the flaps progressively moved o_tboard as the wing tilt angle was increased. For a wing tilt of 40 ° an_ flap deflection of 60 ° and !

T c values 3.5 to 7.4 the air flow was attach+d _et_+rt_e surface of all the vanes and flap except the flap segment outboa_'_ ,_f the outboard nacelle.

(Masted tufts a few inches above the flap surfac_=_+ indicated "separated" flow even when surface tufts i_dicated attached flow.)

Effect of Wing Leading-Edge Devices

Several high-lift devices were tested in various partial- and full-span

configurations on the wing leading edge. Air-flow separation from the out-

board portion of the wing between the nacelles was delayed when the partial-

span KrUger flaps were replaced with partial-span slats. The onset of

air-flow separation over the outboard portion of the wing and the wing root

was further delayed when the spanwise extent of the slats was increased to

cover the exposed wing leading edge from the fuselage to the wing tip.

Longitudinal characteristics for the full-span slat configuration are

presented in figures 14 and 15 for wing tilt angles of 20° and 30° and flap

deflections of 60°• Descent angles, based on CLmax , of 4° to 6° were

obtained with the full-span slats. The angle of attack at which air-flow

separation occurred outboard on the wing was significantly increased (no such

separation occurred for any of the test points shown in fig. 14 or 15 at

T$ _ 0). However, this large delay in separation was not reflected in the

descent performance, based on CLmax,because the separation at the wing

center section largely determined the angle of attack for maximum lift. The

buffeting characteristics of the horizontal tail remained about the sameas

those of the basic configuration.

Several wing center section configurations were tested to determine the

extent to which the air-flow separation could be delayed or contained. The

configurations included various combinations of the basic short fore and aft

ramps, modified fore and aft ramps about twice the length of the basic ramps,

a slat replacing the fore ramp, and fences (extending from the front edge of

the fore ramp to the aft edge of the aft ramp) on the wing above the sides of

the fuselage. A comparison of the force data is presented in figure 16 for a

configuration with the basic short fore and aft ramps and for a configuration

with a long fore ramp with added side wrap-around fairings and a long aft

ramp. None of the wing center section configurations tested significantly

reduced the air-flow separation problem or changedthe buffeting character-

istics of the horizontal tail. Figure 16 showsthat no improvement in force

data was obtained.

For a tilt-wing aircraft with an exposed wing center section there

appears to be little point in optimizing leading-edge devices for increased

descent performance until separation has been eliminated from the wing center

section.

Effect of Spanwise Distribution of Propeller Thrust

Tests were conducted to determine the effect of varying the wing span-

wise lift and drag distribution by operating the inboard and outboard pro-

pellers at different thrust settings. For each run_ the inboard and out-

board propeller speeds were equal and held constant with the inboard propeller

blade angle higher than the outboard. The propeller thrust was controlled by

propeller blade pitch angle in an attempt to obtain higher downwash angles

in the slipstream due to the downward rotation of the propeller over the wing leading edges in the region of the fuselage where the effects of air-flow separation were predominant in determining maxLm_ lift coefficient. The descent performance should then be improved in two ways: (i) by the higher drag at a given lift coefficient obtained by a distorted span loading and; (2) by the increase in the maximum values of lift coefficient before the onset of air-flow separation.

Several tests at various differential propeller blade angle settings are compared at the same average thrust coefficient in figure 17 for a 20 ° wing- tilt configuration. Both maximum lift and descen_ capability increased as differential propeller thrust was increased.

The largest differential propeller thrust settings tested were with an inboard and outboard propeller blade angle of 14 ° and 0°_ respectively.

Force data at several average thrust settings are shown in figures 18 and 19 for the _ = 14/0 configuration with wing tilt an!:_ies of 20 ° and 30_ flaps deflected 60 ° , and full-span slats. It is seen that the descent capability was significantly increased to descent angles greater than i0 °.

The individual propeller thrust coefficierrts shown in figures 18 and 19 are approxJs_ate valuos obtained from two thrust calibrations_ one with differ- ential blade angle settings of 14 ° and 0 ° and the other with all propellers set at 14 ° (see fig. 0). At a given advance ratio_ the thrust of the inboard propeller set at } = 14 was therefore taken to bo independent of the thrust setting on the outboard propeller. With this ass<a_ption the thrust of the outboard propeller with _ : 0 varied from s]i&gtb _ negative to slightly positive for the range <of average thrust coeri'icien'_s sno_m on figures i_ and 19.

With full-span slats on the wing, and average thrust coefficients of i.i or more_ the flow over the portion of the wing behind the outboard propeller did not separate until the geometric wing angqe of attack was 40 ° or more.

With the outboard propellers removed; the outboard portion of the wing sepa- rated at a wing geometric angle of attack of _ = .

o_o (T c 2 6) and as shown on figure 17, the two-propeller configuration had _._ssontially no descent capa- bility. Why an outboard propeller producing little positive thrust if any_ protects the wing much more than no outboard propeller is of interest and worthy of future investigations. A strip analysis of the propeller set at = O_ for the conditions of figure 17_ showed that approximately the inboard half radius of the highly twisted blades was producing positive thrust although the integrated propeller thrust was sli}<htly negative. Since the tip of the outboard propeller overlapped the i_i:} _'d propeller and extended beyond the wing tip, a significant positive pr{}ye]ler slipstream component probably existed over most of the outboard portion of the wing even if the outboard propeller was producing a little negative total thrust.

The differential propeller thrust operation affected the span loading so that the drag at a given lift coefficient was greater, and it also delayed the onset of wing center section air-flow separation. With the wing tilted .... O 7 ° __0 °, aft l'usola_e air-flow separation was delayed for about _ to fuselage an_!u of attsx_k_ which was a gro<_ter improvement than that realized from any of the modified ramp, slat_ or fence combinations tested. TI_e decrease i_ <k, _ k,'_rizont:.i]-t<<ilbuffet inteL_sities as a result of delayin<< the wing center s<,,,:tio__ :Ji!:-i'l<x,, _ separa_tion is sho}rn in figure 12.

Sw!mmz'y ibrce data comparing the differential propeller thrust confi urs_- tion to the !'_artial-span KrUger flap and full-span slat configurations provi- ous!y discussed are sho}_ in figure 20.

Descent Performance Descent performance is summarized in figure 21. For the subject tilt- wing model, descent performance depends upon vSat one considers the i_iting condition to be, that is, wing center section air-flow separation, aft fuse- lace air-flow separation, wing root air-flow separation_ a particular horizontal-tail buffet intensity, CLmax, the angle of attack at which air-flow separation occurs on the outboard portion of the wing_ and so forth.

If the first appearance of wing center section air-flow separation were selected as the criteria, it could be shown that the subject model in any configuration tested would not be capable of i g level flight throughout the transition range. This would probably be a pessimistic interpretation of the tests. Figure 21 is based on the angle of attack at which wing air-flow separation occurred outboard of the inboard nace!le_ which in many cases was beyond the angle of attack for CLmax" This presentation is optimistic because no flight margin is included_ the fuselage angle of attack is not restricted, and the other possible limitations previously mentioned are ignored. Thus figure 21 represents a maxim_mm possible descent performance boundary. The degree of optimism of figure 21 can only be determined by flight testing the confisuration.

Figure 21 shows that for the basic configuration (partial-span KrUger flaps and trailing-edge flaps deflected 60 °) the descent capability is 5 ° or less at airspeeds below 60 knots. Descent performance, based on outboard wing air-flow separation_ was significantly increased by incorporating full- span slats and by the use of differential propeller thrust. With the wing tilted 30 ° and for an airspeed of 50 knots; rate of descent was about 200, 700, and 1200 ft/min, respectively, for the configurations with partial-span Kr[ger flaps, full-span slats, and full-span slats with differential propeller thrust.

Longitudinal Stability and Control Figure 22 presents the longitudinal static stability (Cm_) at various thrust coefficients and wing tilt angles for the basic configuration witk and without a tail and with the fuselage at 0 ° angle of attack. With the tail oYf

and no wing tllt, Cm_ was slightly positive (tm:stable) and becamemore unsta-

ble as either wing tilt angle or thrust coefficient increased. With or with-

out the tail the stability characteristics were s:-milar, but without the

tail, the Om_ was more unstable. At higher rusc_!age angles of attack, tail

on or off, C_ values were similar to those of ficure 22 or somewhatmore

u_us tab le.

The effectiveness of the horizontal tail l;_ presented in figure _:j_i'_r

wing tilt angles of 0°, 20° , and 40° and flab, s £e:'lected 60° •

Figure 23 showsthat the pitching mor:_ent wa_ nearly trirm_:<.d with the

_ T' of 2.0, and the incidence of the horizontal tail set

wing tilted 90°, a -c

at 0°. This point corresponds to a shallow desceit at an airspeed of 30 knots

(W/S = 70). It is also seen that pitch control with only tail incidence was

not sufficient to trim the 20 ° wing tilt confic,__r%tion for a T c of 3.5, which corresponds to shallow climbing ±'light at a:.out 55 knots airspeed.

These results are %_ical throughout the transition range tested; that is th:: model could be trinmed with only tail incidence f:_r i g level flight for win_ tilt angles from 0 ° to 400 at the low foz_,zard ,:peeds, but an additional pitch control device_ such as a tail rotor, would %o r_iuired for climbing flight.

The largest measured value of Cmit was ,_. C_' -P per degree_ obtained with the wing untilted and the propellers operat:i_r :_,,t i_ifferentiai thrus_ settir_:: such that the inboard propeller T_. was about 9.0. Values of Cmit oz' 0.0;:> to 0.040 were obtained with the wing tilted 20 ° t,: 40 °. This variatior, i:: Cmit was probably caused by variations i_! ti_e q_/c_:, ratio as descrR_ g i_.

the following paragraphs.

The fuselage angle-of-attack range for which the inboard propeller slip- stream significantly increased the velocity over tke tip of the horizontal tail is shown in figure 24 for the wing at 0 _ tilt angle and the trailing- edge flaps deflected 60 °. The outboard qt/',_ me,_sured near the tip of the horizontal tail, reached a maxim'_ of 4.3 at _: = e5 °. The inboard _t/;_-, measured at a spanwise location above the si<}_ <_i _ the fuselage, was ap_-roxi- mate!y 0.7 throughout the angle-of-attack rampu.

For wing tilt angles of 20 ° and 30 ° and i'!ap deflections oi' 60 °, the outboard i'ai_:idlyincreased and/'or (]er'r_.ar_:d as _ was increase( b depending on the thrust setting and separation _;h_racteristics over the win_ _ center sections and outboard portions of the win_ (fig. 25). Inboard values of 0 were obtained for wing tilt a_les or 20 ° or 30 ° at the hi£:er fuselage angles of attack as typically showr: ],;,j J'igure 25.

The horizontal-tail buffeting was a i'un'_:_-::ion ,of the air-flow sq_ar-_tiox over the wing cente; _ section and the effec:t ol' fix inboard propeller s!ip- stress_ on the local velocity over the tip o_' tiu burizonta! tail. Lore< Luf_'_ t occurred either whenever the wake from the _, _,'x_,ted center _ection entre!g:0 .i the tail or whenever the outboard r%t:i_:, _.':._ .much greater that_ tk_ inboard qt/_,_ due to slipstream effects.

Comparison of Theoretical and Experimental Wing Stall Boundaries

A comparison of theoretical and experimental wing angle of attack for

CLmax and the wing angle of attack for the onset of wing air-flow separation

outboard of the inboard nacelle is shown in figure 26. The comparison is

of slipstream to free-stream velocity. The theory (ref. 4) is a frequently

used one in which the free-stream and propeller componentvelocities are

vectorially added for estimated wing stall boundaries for tilt-wing VTOL

aircraft. The theoretical assumptions are that (I) the slipstream component

of an inclined propeller can be computed from simple momentum theory, (2) the

wing is fully immersedin a fully developed slipstream, and (3) rotational

effects of the slipstreams are small.

At (T/A)/q_ of 4, experimental CLmax occurred at about 35° wing angle

of attack comparedto a predicted value of 65° angle of attack. Oneof the

reasons for the lack of correlation is that CLmax was limited by air-flow

separation at the wing center section instead of directly behind the propel-

lers. It is apparent that predicted descent performance of a tilt-wing air-

craft may not be realistic if areas that are not immersed in the propeller

slipstreams are ignored. Adjusting the theory to account for only the speci-

fic area outside the propeller slipstreams may also yield optimistic estimates

since air-flow separation which originates in the unprotected area may pro-

gress into significant portions of the immersedarea.

The theoretical boundary for the onset of wing air-flow separation was

also optimistic for this model when the area of interest was limited to the

portion of the wing outboard of the inboard nacelle which was immersed in

the propeller slipstream. Air-flow separation originated in this outboard

wing area, instead of being the result of spreading from an unprotected area

into the immersedarea. At a (T/A)/q_ of 4, air-flow separation occurred

between the nacelles at about 57° wing angle of attack comparedto a pre-

dicted angle of attack of 80°. The theory may be empirically corrected to

agree with the experimental angle of attack for outboard wing air-flow

separation on this model if the propeller thrust is assumedto be about 0.6

of the actual thrust.

The above discussion applies to the typical operation in which all four

propellers produce approximately equal thrust. As stated in the section on

differential spanwise propeller thrust, separation over the outboard portion

of the wing was delayed for an additional 12° angle of attack with the out-

board propellers at approximately zero thrust comparedto the results with

the outboard propellers removed. Thus, for the sameaverage conditions, local conditions signficantly influenced experimental wing stall boundaries.

ii

CONCLUSIONS

Wind-tunnel tests of a large-scale tilt-wing model indicate that signi-

ficant improvement can be madein the stall boundaries and descent performance

in the low-speed transition speed range. For a wing tilt of 30° at 50 knots

the descent performance was increased by 500 ft/min when full-span slats were

used instead of partial-span KrUger flaps and an additional 500 ft/min

increase in descent performance can be obtained by differential propeller

thrust across the wing span.

Theoretical boundaries for air-flow separation or stall were grossly

optimistic comparedto the experimental results obtained for the model of

this investi6ation.

AmesResearch Center

National Aeronautics and SpaceA_ninistration

Hoffett Field, Calif., July 17, i)64

REFERENCES l° Weiberg, Jss_es A., and Holzhauser, Curt A.: Large-Scale Wind-Tunnel Tests of an Airplane Hodel With an Unswept, Tilt Wing of Aspect Ratio 5-5, and With Four Propellers and Blowing F!a!Ds. NASA TN D-I034, 1961.

. Weiberg_ James A., and Giulianetti, Demo J.: Large-Scale Wind-Tunnel Tests of an Airplane Hodel With an Unswept Tilt Wing of Aspect Ratio 5.5, and With Various Stall Control Devices. NASA TN D-2133, 1764.

Pegg_ Robert J.: Summary of Flight-Test Results of the VZ-2 Tilt-Wing _o Aircraft. NASA _{ D-989, 1962.

o Cromwell, Charles H., III, and Payne, Henry E._ III: A Stability Analysis of Tilt-Wing Aircraft (Analytical). Princeton University Rep. 477_ 1960.

TABLEI.- GEOMETRIC DIMENSIONS OFTHEMODEL

Horizontal Vertical

Dimens ion

Wing

surface surface 196.5 50.4 46.7

Area_ sq ft

16.0 40.5

Span, ft 9.35

c, ft 4.99 3.27 5.60

8.35

Aspect ratio 5.08 i. 87

Taper ratio .55

.50 .25 3.7 ° Geometric twist, deg Washout 0 Dihedral from reference 0 0 -2.12 plane_ deg Modified 0018 Root 0015 Root Airfoil section NACA 230_ 0012 Tip oo12 Tip 6.67 Sweep of leading edge, deg 14.7 32.7 Sweep of c/4, deg ii. 0 4.7 25 -7 0 0 Sweep of trailing edge_ deg -1.3 6.26 4.2o 8. oo Root chord, ft 2.10 Tip chord_ ft 3.44 2.00 TABLE II.- STREAMWISE COORDINATES OF WING, _ AND VANE IN PERCENT OF WING CHORD Vane Wing J- Flap X X Yiowcr Ylower Yupper x Yupper[Ylower Yupper 1.41 -2. O0 0 0 0 0 2,72 -1.19 -1.1) 2.82 .8 .61 -2,85 -2.4 b 3 78 • 39 •96 -.96 4 6o -3.49 1,45 -2.8!, 4,23 1,7 ,77 1,32 -1.20 5,64 2.06 5 29 -3.98 2,5 -2. R5 1,15 1.59 -1,36 2.58 7. o5 5.96 -4.37 -2.95 1.54 1.78 -1,45 3"3 lO.58 7.32 -5.15 5,0 3.45 -2.7! 2,31 2.16 -1.53 8.40 6,6 14.11 -5.64 4,05 -2.50 3,08 2,40 -i. 46 8,2 17.64 9.22 -5._ 4,43 -2.1_ 3.85 2.60 -1.27 --6 OI 21.17 9,62 .___ 9.9 4,50 -1.94 4,62 2,75 -.98 f -1.6d 28.22 9.85 -0.70 11.5 4,35 5,39 2.82 -,67 f j 4. oz -i.42 35.27 9,70 -,5.6) o. 1o 2.86 -- 35 i3.2 42,33 -6.7o 14,85 3.68 6.93 2.87 -,07 9,30 -i. 17 8,62 49,38 -6,35 i6.50 -. 96 7,70 2.86 .15 3-33 -5.86 56.42 7,77 18.15 3.02 8.46 2,80 ,31 -- • 7 i,?

-.60 i9.80 63,45 6.87 -5.25 2.67 9.24 2,69 •45 2. Ol 7O.5O 5.78 -4,47 23, i0 -.31 10,77 2,33 ,59 4.61 26,40 -._ 77,60 -3.56 1.35 !.2.32 1.79 ,54 84.6o .69 0 3,30 -2.57 29,70 13.85 •99 •35 i.86 14,62 ,52 ,20 91,70 -1,52 31,40 .31 -.03 i00. O0 .O4 0 • 17 -.17 33-0 25.40 . o7 o m.m.

radius = 2.1_o c L,E, radius = 0.56% c L.E. radius = 0.215 e 123017 airfoil with modified leading edge QJ -p -p q-_ I cO ,-Q I _Q (1) ..p or-I H 11) nd (1) taO I © rio orl z5

r_

\

\ rd

:1°

"-I_ °

,_ .4 ® 20 m .2 I0 0 0 .2 .4 .6 .8 1.0 Fraction of blade radius Figure 7.- Propeller blade characteristics.

I0 I Tc iti i: '.

.I .2 .5 .4 .5 .6 .7 .8 .9 J or CTs Figure 8.- Propeller thrust chara<_tcristics.

5w = 00 WL 33.5 --i I t I WL 31.5 158.4 156.3

I

Moment center WL 13.1 e

I

Fuselage sta. 155.3 WL Fuselage WL 34.1 20 ° WL 31.5 .25c. c_ I 157.1 e WL 14.7

I

Sta. 155.2 WL 0 WL 34.4 WL 31.5 158.

158.4

I

40 ° e WL 16.1

I

Sta.

155.o WL 0 Figure 9.- Variation of pitching-moment center with wing tilt angle.

o It 7) OJ -g -,o ° O0 •+d 0 -P Ca _4 •rd "r-I CH I 0::1 ,D C._ 0 h b_ OH 0 LO CH p__ I CO rd co ._ cg © L> c_ 0 a d 'L} oO o,--I I .rd I t d :2) 0 cO / .r-t © 2_ Od -H +_ bO +_ "H .H bO -H q_ ,% • ©O cH O LD •H IT _3 ,'X) o3 © bl .M © © O -@ .r-I ol ® 4J .r-I 4_ I c_ tt ed?

.r-t O bl?

N

O _3 ® ,-'t

N

O O o II O Od o +-+-_--_- O II t_ o J -p O b_ .H .H O iiii _±iii q_ I1 I ],_H- IIII iiiii !r!r! hi?

O ::::: qd ::::; O [111 1 [1[I I lllll hi?

II_l 1 Illl lili]l • 1 iii i !!!!_ O .H !,]11[ ::::: _LI i i [ e3 11111 iiiii 11,111 +_ I!,111 iiiii !l_ll iiii!

j i II, i i O illll o iilii Ii IIII IliJ I I I Ill !!!!!

IIIII IIIIli r_ I I I I i'll !!!!!

".,,%'I', Od .H +-_ -++.+ +_ ::÷_+ C) (1) iiiii -H_ O !!!!_ ;:::; iiiii !

iiiii E II11l r-t lltlt iiiii I'-.- CO If) _ rO (Xl O b.0 ,r-t 0 = _ 7:::) _ 7 3 = 7:::) V r_ O0 ..J (D I _3 C_ I © O © H q_ ;O © O 0"% C3 © .r-I +-_ t_ I .rq o _0 Od 4° ii I ®o O o o] t,o I © [] _D o +_ +) rT] OO @ +_ C) O CO _c_ I +) _D I I r-J © Od 0 I bO °_ o oJ o 3o ..p _3 r_ P_ r_ i ,-t o o IF q4 °_ o O O.l r+d o -p ,-t °_ °r-I egO a O O r.j _0J -p • H rl -p LQ .H © ,--I b:.0 aS I rd ,--I "r--t -p I (b Q q-I rH ",--I "d rH -p O ® rH rH I © .r-t c I , i O O O O _',,j t_ o _D 4s .4D :_ .H Z:' 'I 4s ,:?1 O @ +_ t,t, ,L) ,--I "H :1) o ,s ;S' s_ 4_ bs I i :D O 00 rj) I rt qq o © '\D II r+q o o o -p ,r-I -p b,g 'H • _3C_ +_ •r-t +._ N._ rt ©.__- x_ O _ -O .rq % o _3 h _3 -M +_ °H o I & © % o ._1 rO q- ...................... ! " -_ 000 000 .................................... _ G'_ I"--- 0,.I o,.10 "_l._ PO bO @ Od o dD "H © .M I Od b_ I ",--I q_ © ....... i ..........

:!!ii

-0 i!

% CO 'i .... l .........

+-_ e3 O0 ......... }!i .

% q_ 0 a ©

)

CO co I !

d (-_ co {1)

i iiiiiii !iiiii i

b_ ,H p... CO LD _- r'O I 0 O3 O0 C.)

3_ O7 O6 .05 .04 q .03 C

.o2

¢J .01 0 2 4 6 8 Tc ' Figure 22.- Lon_:itudimal static stability ,_ith pa_tia!-spam Kr'dger flaps at . . . O

O

C_ -I -2 -3 -I0 0 I0 20 30 40 it , deg Figure 23.- Effectiveness of horizontal tail with short_ fore and aft center- section ramps at various wing tilt angles; Sf = 60 ° .

4.4 32 40 Figure 24.- Dynamic-pressure variation at the horizontal tail; wing tilt of 0°, _f = 60 °, clean leading edge_ _ = IL/0, it = 20 ° .

I I Horizontal toili g 4-- cr -16 -8 0 8 16 24 Figure 25.- Dynamic-pressure variation at the horizontal tail; wing tilt of 20o3 bf = 60o3 leading-edge slat outboard of inboard nacelle 3 = i0 o, i t = 20 o.

9O

8O

7O

6O

5O

_W 4O 5O a for CLmax , limited by wing center section separation i i 2O I0 0 I 2 5 4 5 6 7 ( T/A)/q Figure 26. - Comparison of theoretical and _x_._rim_ntal_ angle of attack for wing stall; leading-edge full-span s ats_ trailing-edge flaps deflected 60 ° • 40 NASA-Langley, 1964 A-8t9 "The aeronautical and space activities of the United States shall be conducted so as to contribute . . . to the expansion o] human knowl- edge o[ phenomena in the atmosphere and space. The Administration shall provide for the widest practicable and appropriate dissemination o[ in[ormation concerning its activities and the results thereof."

--NATIONAL AERONAUTICS AND SPACE ACT OF 1958

NASA SCIENTIFIC AND TECHNICAL PUBLICATIONS

TECHNICAL REPORTS: Scientific and technical information considered important, complete, and a lasting contribution to existing knowledge.

TECHNICAL NOTES: Information less broad in scope but nevertheless of importance as a contribution to existing knowledge.

TECHNICAL MEMORANDUMS: Information receiving limited distri- bution because of preliminary data, security classification, or other reasons.

CONTRACTOR REPORTS: Technical information generated in con- nection with a NASA contract or grant and released under NASA auspices.

TECHNICAL TRANSLATIONS: Information published in a foreign language considered to merit NASA distribution in English.

TECHNICAL REPRINTS: Information derived from NASA activities and initially published in the form of journal articles.

SPECIAL PUBLICATIONS: Information derived from or of value to NASA activities but not necessarily reporting the results of individual NASA-programmed scientific efforts. Publications include conference proceedings, monographs, data compilations, handbooks, sourcebooks, and special bibliographies.

Details on the availability of these publications may be obtained from: SCIENTIFIC AND TECHNICAL INFORMATION DIVISION NATIONAL AERONAUTICS AND SPACE ADMINISTRATION Wadlinston, D.C. 20546

Overview