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NASA Technical Memorandum 10/,083 . Experimental Aerodynamic Characteristics of a Joined-Wing Research Aircraft Configuration Stephen C. Smith and Ronald K. Stonum April 1989 (NASA-TM-|0 I083) EXPERIMENTAL AERODYNAMIC CHARACTERISTICS OF A JOINED-WING RESEARCH AIRCRAFT CONFIGURATION |NASa. ames Research Center) 100 p CSCL 01A G]/02 N89-2_265 Unclas 0217216 NASA National Aeronautics and Space Administration https://ntrs.nasa.gov/search.jsp?R=19890014914 2020-05-24T21:39:48+00:00Z

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Page 1: Experimental Aerodynamic Characteristics of a Joined-Wing … · 2014-10-04 · Experimental Aerodynamic Characteristics of a Joined-Wing Research Aircraft Configuration Stephen C

NASA Technical Memorandum 10/,083 .

Experimental AerodynamicCharacteristics of a Joined-WingResearch Aircraft Configuration

Stephen C. Smith and Ronald K. Stonum

April 1989

(NASA-TM-|0 I083) EXPERIMENTAL AERODYNAMICCHARACTERISTICS OF A JOINED-WING RESEARCH

AIRCRAFT CONFIGURATION |NASa. amesResearch Center) 100 p CSCL 01A

G]/02

N89-2_265

Unclas

0217216

NASANational Aeronautics andSpace Administration

https://ntrs.nasa.gov/search.jsp?R=19890014914 2020-05-24T21:39:48+00:00Z

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NASA Technical Memorandum 101083

Experimental AerodynamicCharacteristics of a Joined-WingResearch Aircraft Configuration

Stephen C. Smith, Ames Research Center, Moffett Field, CaliforniaRonald K. Stonum, Air Force Systems Command Liaison Office, U. S. Air Force,Ames Research Center, Moffett Field, California

April 1989

NASANational Aeronautics and

Space Administration

Ames Research CenterMoffett Reid, California 94035

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NOMENCLATURE

b

c

c/4

CD

CL

Ct

Ct

era

C,,

C_

Ctn

Ct,

C t,,

OCt�OCt.

C_FI

FM

FO

L/D

L.E.

RI

RO

S

U_

15

Xt.9.

wing spanchord

25% chord line

centerhne

coefficient

coefficient

coefficient

coefficient

coefficient

coefficient

of dragof lift

of lift of airfoil section

of rolling moment

of pitching moment

of yawing momentcoefficient of side force

lateral stability--change in rolling moment with sideslip

roll control power--change in roiling moment with aileron deflection

roll damping--change in rolling moment with roll rate

longitudinal stability--change in pitching moment with lift

directional stability--change in yawing moment with sideslip

forward wing inboard control surface

forward wing midspan control surface

forward wing outboard control surface

lift to drag ratio

leading edge

rear wing inboard control surface

rear wing outboard control surface

wing area

aircraft flight speed

minimum nosewheel-raising speed

power-off stall speed

distance from center of gravity to main landing gear

Greek Symbols

a angle of attack, deg

angle of sideshp, deg

control surface deflection, deg

¢ angle of roll, deg

roll rate, rad/sec

Subscripts

a

c.g.

max

r

ref

aileron

center of gravitymaximum

rudder

reference

°°,

in PRECEDING PAGE BLANK NOT FILMED

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SUMMARY

A wind-tunnel test was conducted at Ames Research Center to measure the aerody-

namic characteristics of a joined-wing research aircraft (JWRA). This aircraft was designed

to utilize the fuselage and engines of the existing NASA AD-1 aircraft. The JWRA was de-

signed to have removable outer wing panels to represent three different configurations with

the interwing joint at different fractions of the wing span. A one-sixth-scale wind-tunnel

model of all three configurations of the JWRA was tested in the Ames 12-foot low-speed

wind tunnel to measure aerodynamic performance, stability, and control characteristics.

This report presents the results of these tests. Longitudinal and lateral-directional

characteristics were measured over an angle of attack range of -7 ° to 14 ° and over an

angle of sideslip range of -5 ° to +2.5 ° at a Math number of 0.35 and a Reynolds number

of 2.2 × 10 e/ft. Various combinations of deflected control surfaces were tested to measure

the effectiveness and impact on stability of several control surface arrangements. In addi-

tion, the effects on stall and post-stall aerodynamic characteristics from small leading-edgedevices called vortilons were measured.

The results of these tests indicate that the JWRA had very good aerodynamic perfor-

mance and acceptable stability and control throughout its flight envelope. The vortilons

produced a profound improvement in the stall and post-stMl characteristics with no mea-

surable effects on cruise performance.

INTRODUCTION

The joined wing research aircraft (JWRA) was designed jointly by NASA Ames Re-

search Center and ACA Industries, Inc. of Torrance, CA as part of a Small Business

Innovation Research (SBIR) program. The purpose of this program was to build a joined-

wing proof-of-concept and research aircraft by modifying the existing NASA AD-1 oblique

wing aircraft. The AD-1 aircraft, shown in figure 1, is described in reference 1. A new

wing system which was representative of a joined-wing, transonic commercial transport

aircraft was designed to fit onto the fuselage of the AD-1, using the existing undercarriage

and fuselage-mounted engines. The design of the JWRA is described in references 2 and 3.

A one-sixth-scale wind-tunnel model of the JWRA was built using the fuselage of

the AD-1 wind-tunnel model. This model was tested in the Ames 12-Foot Pressure Wind

Tunnel to measure aerodynamic performance, stability, and control characteristics. This

report presents the results of these wind tunnel tests.

TEST FACILITY

The testswere conducted in the Ames 12-Foot Pressure Wind Tunnel. Figure 2 shows

the JWRA model installedin the wind tunnel test section. The model was supported on

a swept blade-type strut mounted on a bipod support system. The blade was swept back

50° from the normal axis of the model and was contoured with a 15%-thick low speed

airfoil.Figure 3 shows the strut geometry.

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MODEL DESCRIPTION

A one-sixth-scale model of the JWRA was built based on the one-sixth scale AD-1

wind tunnel model. The model had removable outer wing panels to represent three different

wing configurations. With the full-sized outer wing-panels installed, the interwing joint

was located at 60% of the semispan.

A three-view drawing of this configuration, called the JW-1, is shown in figure 4. The

JW-2 configuration had shorter outer panels so that the joint was located at 80% of the

semispan. A three-view drawing of the JW-2 is shown in figure 5. The JW-3 configuration

had the outer panels removed so that the interwing joint was located at 100% of the

semispan. The JW-3 configuration is shown in figure 6. Detailed dimensional data on all

three configurations ares given in table 1.

The design of the airfoils for the JWRA is discussed in reference 2. The wing geometry

was produced by locating each airfoil perpendicular to the 25% chord-line and lofting

straight lines between these defining sections. The coordinates of the airfoils are given in

table 2.

The JWRA wings were fitted with five control surfaces on each side, representing

various arrangements of flaps, elevators, and ailerons. The nomenclature for the various

control surfaces on the wings is shown in figure 7. The vertical tail was fitted with a

conventional rudder. The control surface deflections available for the model are given in

table 3. The sign convention for control surface deflections was chosen to be trailing-edge

down as positive for all the wing surfaces and trailing-edge right as positive for the rudder.

The JWRA wings were designed so that small undersurface leading-edge fences called

vortilons could be mounted at three span stations. A detail of the vortilon geometry and

the three mounting stations is shown in figure 8.

The engine pylons were designed so that speedbrakes could be installed between the

fuselage and nacelles. A detail of the speedbrake geometry and installation is shown in fig-

ure 9. The model had simple flowthrough nacelles without instrumentation for measuring

internal duct losses.

TEST CONDITIONS AND PROCEDURES

At the time these tests were conducted, the 12-Foot Pressure Wind Tunnel could not

be pressurized. Therefore, the tests were conducted with the wind-tunnel total pressure

equal to the ambient atmospheric pressure. Model strength limitations established a max-

imum dynamic pressure of 170 psf. Almost all the data from these tests were taken at

these conditions, which produced a test Mach number of 0.35, a test Reynolds number

of approximately 2.2 × 106/ft, and a chord Reynolds number of approximately 1.0 × 106.

It was found that some of the deflected aileron configurations would have exceeded the

rolling-moment capacity of the force balance, so it was necessary to reduce the dynamic

pressure to 120 psf for some tests, and in a few cases, to 90 psf.

The angle of attack was varied from -7 ° to 14 ° and the angle of sideslip was var-

ied from -5 ° to +2.5 ° . The angle of attack was measured by an electronic inclinometer

installed in the model directly below the force balance. The angle of sideslip was men-

2

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sured from the turntable portion of the wind-tunnel model-support system. The measured

sideslip angle was corrected for model support and balance deflection.

A 2.0 in. internal six-component strain-gage force balance was used to measure aerody-

namic forces and moments. The moment reference center was located at the assumed c.g.

location for each configuration, which was based on a static stability margin of 0.35c.el.

The reference area for computing aerodynamic coefficients was the projected plan area of

the forward wing. The moment reference center and reference area for each configuration

are given in table 1. The axis system and sign conventions for aerodynamic data in this

report are shown in figure 10. Lift and drag coefficient data are presented in the wind axis

system. All other aerodynamic coefficients are presented in the body axis system.

The data presented in this report have been adjusted for wall effects and blockage

effects. The wall corrections were determined by the method described in reference 4,

and the blockage corrections were determined by the method described in reference 5. No

adjustments were made to remove internal drag from the nacelles or the small cavity in

the fuselage where the blade support attached to the balance.

RESULTS AND DISCUSSION

Experimental aerodynamic data for the JWRA are shown in figures 11 through 84.

The results for the JW-1 configuration are shown in figures 11 through 41. The JW-2

results are shown in figures 42 through 58. The JW-3 results are shown in figures 59

through 75. Results for the JW-1 wing-body (rear wing removed) are shown in figures 76

through 79, and results for the body only are shown in figures 80 through 84.

JW-1 Configuration

Longitudinal characteristics (zero control deflections)- The longitudinal characteristics

of the JW-1 are shown in figures 11 through 14. Figures 11 through 13 show the variation of

lift, drag, and pitching moment with angle of attack. The variation of longitudinal stability,

(OCm/OCL), with angle of attack is shown in figure 14. A maximum lift coefficient (CL,,,,.)

of 1.20 was obtained at a 6* angle of attack. This is a satisfactory value for a wing with

30°of sweep and maximum section lift coefficent (Ct,,,..) of 1.5. The apparently low angle

of attack for stall is the result of the large wing incidence relative to the body axis (see

table 1) which was necessitated by the short landing gear of the AD-1, as discussed in

reference 2. The cruise lift-coefficent of 0.45 was achieved at a -3 ° angle of attack.

There was a reduction in lift curve slope between 2_ ° and 6 ° angle of attack which

indicated a growing region of flow separation on the wing. Figures 13 and 14 show a loss of

stability associated with this separation region. The JW-1 exhibited classical swept-wing

stall characteristics where the aircraft was unstable at the stall angle, and became stable

again at a high, post-stall angle of attack of, in this case, about 12 °. However, as shown in

figure 13, this unstable stall was not severe, since a nose-down pitching moment still existed

with neutral elevator control and nose-down elevator control would easily reduce the angle

of attack below stall. Figure 14 shows that the JW-1 had a static stability margin of 0.30

to 0.36 over the normal flight envelope (CL=0.3-1.0) for this moment reference location.

3

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Lateral-directional characteristics- The lateral-directional characteristics of the base-

line JW-1 are shown in figures 15 through 19. Figures 15 through 17 show variations in

side-force, yawing-moment, and rolling-moment coe_cients with angle of attack for various

angles of sideslip, B = 0.0 °, +2.5 °, and - 5.0 °. There were abrupt changes in the lateral-

directional characteristics above the angle of attack where the wing stalled. These changes

are usually associated with asymmetric separations due to minor geometric asymmetries

in the model. The untrimmed rolling moment, yawing moment, and side-force diminish

at higher angles of attack, so no unusual or violent departure characteristics are expected.

Figures 18 and 19 show variations in lateral stability, Ct_, and directional stability, C,_,

with angle of attack. Figure 18 shows an unusual loss of lateral stability at high angles

of attack. It appears that the positive dihedral contribution from the forward wing was

reduced as the wing stalled and the negative dihedral of the rear wing began to dominate.

Although the reduction of lateral stability at high angles of attack would have some effect

on the post-stall flying qualities, it is not expected that this would be serious, especially

considering the available lateral control, as described below.

Forced transition- The model Reynolds number for these tests was about one-third of

the flight-vehicle Reynolds number. Transition strips were installed on the model to study

the effects of forced transition on the JWRA model. The primary concern was that the

low chord Reynolds number might lead to areas of premature laminar separation on the

model that would not occur on the full scale aircraft. This would not only affect the drag

data, but also the lift and pitching moment data.

Transition strips consisting of a random distribution of 0.009- to 0.010-in. glass spheres

were installed on the model. The particle size was selected based on the method of refer-

ence 6. Two fixed transition locations were studied; figures 20 through 22 show the results

of this study. First, transition was fixed at 35% chord on all the wing and tail surfaces

and on the fuselage just ahead of the crest of the canopy. This was where natural laminar

transition was expected to occur on the full scale aircraft. As shown in figures 20 through

22, use of this transition location caused some increase in the drag and also affected the

lift and pitching moment characteristics slightly. If a premature laminar separation was

causing loss of stability, the transition strip would be expected to help the problem by

delaying separation and increasing Cl.,,.,. However, the use of the transition strip seemed

to aggravate the problem of stability loss rather than lessen it. To verify that no laminar

separation was occurring forward of the transition strip, a second test was done with the

transition strip installed at 10% chord on the wing and taft surfaces. Figure 20 shows no

further increase in the drag from fixing transition at 10% chord. Based on these results,

it was believed that the drag increase was caused by the drag of the transition strip itself

and a reduction of the laminar run on the fuselage.

The transition strip at 10°/0 chord was left in place for all the other testing of the JW-1

configuration.

Vortilons. The baseline JW-1 configuration showed undesirable loss of stability re-

suiting from outer wing-panel stall. This is typical of high aspect-ratio swept wings with

moderate taper: the lift distribution becomes more heavily loaded near the tips as the

angle of attack is increased, and the wing chord is reduced at the tips because of taper, so

4

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the higher lift loading results in very high lift coefficients on the outer wing panels. The

outer wing-panels frequently stall first, which shifts the center of lift inboard and forward,

producing undesirable pitchup at stall.One device which has been used to reduce this unstable stall characteristic is a small

undersurface leading-edge fence called a vortilon. The use of vortilons for improving stall

and post-stall flying qualities is discussed in reference 7. The JW-1 was tested with vor-

tilons installed at three different span locations, and combinations of these. Figures 23 and

24 show the effect of each of the three vortilon locations on the lift and pitching-moment

characteristics. Each of the three vortilon locations had some influence on the prestall

separation, and the post-stall lift and moment characteristics.

The inboard vortilon configuration had the most significant'effect on the post-stall

pitching moment. Figures 25 and 26 show more clearly than figures 23 and 24 the impact

of the inboard vortilon installation on the longitudinal characteristics of the JW-1. The

eLm., of the JW-1 was increased from 1.20 to 1.22 and the character of the stall break

changed to a much more gentle stall. As shown in figure 26, the unstable pitchup was

eliminated and the JW-1 with inboard vortilons had nearly constant pitching moment

above stall for this moment reference location. A small forward shift in the c.g. location

would make the stall characteristics completely stable. Figure 27 shows that there is no

cruise-drag penalty for the installation of the vortilons. Figures 28 and 29 show the effect

of the vortilons on the rolling moment and yawing moment. The vortilons delayed the

sudden change in lateral-directional characteristics until well above the stall angle.

Figures 30 and 31 show the effect of two vortilons, installed at the inboard and out-

board locations, compared with the effect of the inboard vortilon only and no vortilons.

The two-vortilon configuration maintained more stability up to the stall, but did not have

as much improvement of the post-stall pitchup as the single inboard vortilon.

Although other vortilon locations were tested, the effect of the inboard vortilons was

so significant that this configuration was used as a new baseline configuration for all of the

remaining testing of the JW-1, JW-2, and JW-3 configurations, except where noted.

Longitudinal control- Two different arrangements of controls were evaluated for lon-

gitudinal control effectiveness. Figure 7 shows the nomenclature used to describe the

individual control surfaces. One control arrangement used both the inboard and outboard

rear surfaces (RI+RO) as elevators for longitudinal control. The second arrangement used

the inboard surface on the forward wing and the inboard rear surface (FI+RI) coupled to

move equal amounts in opposite directions so that nose-up pitch command would cause FI

to move downward as a flap and RI to move upward as an elevator. Data for the FI+RI

control arrangement were determined by interpolating wind-tunnel data from configura-

tions with 6Rr = -15 ° and gFr = 0 ° and 20 °, since model hardware for _SFt = 15 ° was not

available. Figure 32 shows the control effectiveness of the RI+RO elevator combination

and the FI+RI flap-elevator combination compared with the effectiveness of RI only. The

FI+RI configuration had slightly greater control authority. As shown in figure 33, the

FI+RI arrangement had the advantage of changing the lift in the proper sense for the

change in pitching moment, i.e., a nose-up moment was accompanied by an increase in

lift due to the coupled flap deflection. However, this arrangement had the disadvantage

of substantially disturbing the designed incidence distribution of the wing at high angle of

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attacks by deflecting the flap to trim the airplane.

One criterion for assessing longitudinal control authority is the minimum speed re-

quired to lift the nosewheel. The ratio of minimum nosewheel-raising speed to power-off

stall speed was determined from moment equilibrium about the main wheels, neglecting

thrust effects. This ratio was expressed as:

= /

_ W XI'9" CL_=° + c"_sC_c g.... o(1)

where the lift and pitching moment coefficients at a = 0 were based on full control deflec-

tions of 15 °, as shown in figures 31 and 32. This ratio was determined for the RI+RO and

FI+RI configurations and the RI+RO configuration with the forward wing surfaces used

as takeoff flaps (6r/ = 20 ° and _rM = 15°). Values of the minimum nosewheel-raising

speed ratio for the three control configurations are given below.

(RI+RO) (FI+RI) (RI+RO)+Flaps

0.97 o.s9 o.s4

The RI+RO+flaps configuration was the most effective, although all three control schemes

were able to raise the nosewheel below the stall speed. The reader should bear in mind

that these values are based on 15 ° control deflections and more control authority would

be available from greater control deflections.

Control surface deflection also affected the post-stall stability characteristics. Both

the RI+RO and FI+RI configurations appear to be completely stable well above the stall.

Trimmed characteristics- Wind-tunnel data obtained with and without the control

surfaces deflected were used to derive trimmed lift and drag characteristics of each longi-

tudinal control arrangement (neglecting thrust effects). Figure 34 shows the trimmed lift

characteristics of the FI+RI configuration, and the RI+RO configuration with and without

the FI and FM surfaces acting as take-off flaps.

Figure 35 shows the variation of trimmed lift-to-drag ratio (L/D) with lift coefficient

for both control arrangements. The FI+RI configuration had a slightly lower L//D because

of the alteration of wing twist as the controls are deflected.

Lateral-directional control- The JW-1 had a conventional aileron surface on the for-

ward wing outboard of the interwing joint (FO). This aileron may be supplemented with

either the middle forward flap (FM) or the outer surface on the rear wing (RO). Each

aileron surface used differential deflection so that the right aileron was deflected -15 ° and

the left aileron was deflected +10 °. Figure 36 shows the control effectiveness of the out-

board aileron alone and supplemented by each of the other two candidate ailerons. Figure

37 shows the effect of these control surfaces on yawing moment. The FM surface was

considerably more effective as an additional aileron than the RO surface. The forward

wing ailerons produced yaw away from the turn, whereas the rear wing aileron produced

yaw into the turn. In "all cases, the yaw coupling was small compared with the available

directional control from the rudder, as shown in figure 38.

6

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To assessthe control authority of various aileron configurations, an estimate of thesteady-stateroll rate for a full-scaleJW-1 wasmadefrom the measuredaileron effectivenessaud an estimate of the roll damping from a vortex-lattice program. Tile steady-state rollrate wasestimated from

- b (2)( '/r, 2_'._

From figure 36. ('(,,,b,, -- 0.035 at _ = 0 for bFo - -15°/ + 10 °. The roll damping,

(',.,, was esliumted _ }>e -0.B. The JW-1 span was 40 ft. and it was assumed to beflying at 1.50 ft sec. The steady-state roll rate for this configuration was estimated to

be 0.44 rad/sec or 2.5' ,sec. Greater aileron deflection would, of course, lead to a greaterroll-rate.

Control effectiveness at high angles of attack is also an important evaluation criterion.

Figures 36 and ?,,_ iudicate that both the ailerons and ru<ider maintain good control au-

thority above .,tall and were more than adequate to control the abrupt changes in rolling

and yawing munwnt above the stall angle-of-attack shown in figures 28 and 29.

,Speed brake- The JW-1 was tested with sinlple speed brakes deployed from the engine

pylons as shown in figure 8. These speed brakes were designed as a means of altering the

glide slope for landing, because of the high L/D of the JWRA and the high idle-thrust

of the turbojet eugines. Figures 39 through 41 show the lift, drag, and pitching-moment

characteristics of the JW-1 with and without the speed ])rake deployed.

.1W-2 ('onfi_;urat.ion

LongitudiT_ol characteristics- Tile longitudinal characteristics of the JW-2 are shown

in figures 42 through 46. The transition strips were removed from the wings for flow-

visualization experiments on the JW-1 and were not reapplied. The JW-2 and JW-3

configurations were tested without forced transition. Figures 42 through 44 show the

variation of lift. drag, and pitching nmment with angle of attack. Figure 45 shows the

variation of longitudinal stability with angle of attack. Tile JW-2 had a static stability

of 0.30 to 0.34 _,ver the normal flight envelope for this moment reference location. Figure

46 shows the l)itchin_ moment characteristics of the JW-2 with the vortilons removed. As

with the JW-1, the longitudinal characteristics without the vortilon are less satisfactory.

L, tcral-di_'¢ctional characteristics. The lateral-directi¢,nal characteristics of the JW-2

are shown in figures 47 through 51. Figures 47 through 49 sh¢,w variations in side-force,

yawing-moment, and rolling-moment coefficients with angle of attack for sideslip angles of

0 °.and zk 2.5 ° . FiKures 50 and 51 show variations in lateral stability, ('el, and directional

stability, (',+, with an_le of attack.

LoTzgit,dinal cotdvol- Both the RI+RO and FI+RI control arrangements were evalu-

ated on the J W-2. As with the JW-1 configuration, data for the FI+ RI control arrangement

were determined by interpolating wind tunnel data. The control effectiveness of both ar-

rangements is shown in figure 52. Figure 53 shows the effect of these controls on the lift

characteristics of the JW-2. The two control arrangements had nearly the same elevator

effectiveness on the JW-2.

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Trimmed characteristics- Wind-tunnel data with and without the control surfaces

deflected were used to derive trimmed lift and drag characteristics of the JW-2 using

both the RI+RO and FI+RI control configurations. Figure 54 shows the trimmed lift

characteristics and figure 55 shows the variation of trimmed L/D with lift coefficient for

both arrangements.

Lateral-directional control. The conventional aileron surface, FO, was shorter on the

JW-2 since the outer wing-panel was smaller. Figure 56 shows the control effectiveness

of the shorter FO aileron, both alone and supplemented by each of the candidate inboard

ailerons. Figure 57 shows the effect of these controls on yawing moment. Figure 58 shows

the directional control effectiveness of the rudder. As with the JW-1, the FM surface was

more effective as a supplemental aileron than the RO surface.

JW-3 Configuration

Longitudinal characteristics- The longitudinal characteristics of the JW-3 are shown in

figures 59 through 63. Figures 59 through 61 show the variation of lift, drag, and pitching

moment with angle of attack. Figure 62 shows the variation of longitudinal stabihty

with angle of attack. The JW-3 had a static stability of 0.28 to 0.38 over the normal

flight envelope for this moment reference location. Figure 63 shows the pitching moment

characteristics of the JW-3 with the vortilons removed. As with the JW-1 and JW-2, the

longitudinal characteristics without the vortilon are less satisfactory.

Lateral-directional characteristics- The lateral-directional characteristics of the JW-3

are shown in figures 64 through 68. Figures 64 through 66 show variations in side-force,

yawing-moment, and rolling-moment coefficients with angle of attack for sideslip angles of

0°and ± 2.5 °. Figures 67 and 68 show variations in lateral stability, Ctn, and directional

stability, C,_o, with angle of attack.

Longitudinal control- Both the RI+RO and FI+RI control arrangements were evalu-

ated on the JW-3. As with the JW-1 and JW-2 configurations, data for the FI+RI control

arrangement were determined by interpolating wind tunnel data. The control effectiveness

of both arrangements is shown in figure 69. Figure 70 shows the effect of these controls on

the lift characteristics of the JW-3. The FI+RI arrangement was somewhat less effective as

an elevator than the RI+RO arrangement on the JW-3. The FI+RI arrangement became

less effective on the JW-2 and JW-3 configurations because the moment reference center

(at the c.g.) was moved forward to achieve the sasne level of longitudinal stability. This

shortened the moment arm of the additional hft on the forward wing caused by deflection

of the FI surface.

Trimmed characteristics- Wind-tunnel data with and without the control surfaces

deflected were used to derive trimmed lift and drag characteristics of the JW-3 using

both the RI+RO and FI+RI control configurations. Figure 71 shows the trimmed lift

characteristics and figure 72 shows the variation of trimmed L/D with lift coefficient for

both arrangements.

Lateral-directional control- The outer wing-panel was removed on the JW-3, leaving

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only the inboard surfacesavailablefor roll control. The FM and RO surfaceswereevaluatedas ailerons. Figure 73 showsthe control effectivenessof the FM and RO surfacesactingindividually and together. Figure 74showsthe effectof thesecontrols on yawingmoment.The FM surface was considerably more effective as an aileron than the RO surface. As

with the JW-1 and JW-2, the FM surface produced adverse yaw, whereas the RO surface

produced yaw into the turn. Rudder deflection data for the JW-3 was taken only with the

ailerons deflected. Figure 75 shows the rudder effectiveness with the FM aileron deflected.

The yaw coupling was quite small and easily controllable with the rudder in all cases.

JW-1 Wing-Body

The JW-1 configuration was tested with the rear wing removed to obtain longitudinal

aerodynamic characteristics of the wing-body only. The vertical tail and engines were left

on the fuselage, and vortilons were installed at the inboard wing station. The area of the

vertical tail where the rear wing attaches was faired smoothly to match the contours of

the vertical tail. The JW-1 moment reference center and reference area were used. Figures

76 through 78 show the variation of lift, drag, and pitching moment with angle of attack.

Figure 79 shows the variation of pitching moment with angle of attack with the vortilonsremoved.

Body Only

Longitudinal characteristics of the JWRA fuselage with vertical tail and engines in-

stalled are shown in figures 80 through 82. The wing-root area of the fuselage was faired

over smoothly. The JW-1 moment reference center and reference area were used for com-

putation of coefficients. Figure 83 and 84 show the directional characteristics of this

configuration.

CONCLUDING REMARKS

A one-sixth-scale model of a joined-wing research aircraft (JWRA) was tested in the

NASA Ames 12-Foot Pressure Wind Tunnel. The JWRA was designed to have removable

outer wing panels to represent three different configurations with the interwing joint located

at different fractions of the wing span. The aerodynamic characteristics of these three wing

configurations, called JW-1, JW-2, and JW-3, were measured. The JWRA was found to

have acceptable levels of stability and control, especially with vortilons installed. All of

the control surfaces maintained su_cient effectiveness to control the JWRA well above the

stall angle of attack.

Two longitudinal control configurations were evaluated for use on the JWRA. The

FI-RI arrangement was found to be slightly more effective than the RI+RO arrangement

on the JW-1. However, the FI+RI arrangement lost effectiveness as the moment reference

center was moved forward on the JW-2 and JW-3, since the pitching-moment contribution

from an increase in wing hit was reduced. On the JW-3, this arrangement was less effective

than the RI+RO arrangement. The FI+RI arrangement had a slightly lower L/D than

the RI+RO arrangement on all three configurations, because of the alteration of the wing-incidence distribution when the FI surface was deflected.

9

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Two control surfaceswere evaluated for roll control effectiveness to supplement the

aileron on the JW-1 and JW-2 and to act as the primary aileron on the JW-3. In all cases,

the FM surface was more effective as an aileron than the RO surface. None of the aileron

surfaces studied created severe adverse yaw.

A small speed-brake arrangement installed on the engine pylons was found to produce

a significant amount of drag, providing a convenient means of glidepath control for landing.

10

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REFERENCES

1. Sim, A.G.; and Curry, R.E.: Flight Determined Aerodynamic Derivatives of the AD-1

Oblique-Wing Research Airplane. NASA TP-2222, Oct. 1984.

2. Smith, S.C.: Cliff, S.E.; and Kroo, I.M.: The Design of a Joined-Wing Flight Demon-

strator Aircraft. AIAA Paper 87-2930, Sept. 1987.

3. Wolkovitch, J.: Joined Wing Research Airplane Feasibility Study. AIAA Paper

84-2471, Oct. 1984.

4. Sivells, J.C.; and Salmi, R.M.: Jet-Boundary Corrections for Complete and Semispan

Swept Wings in Closed Circular Wind Tunnels. NACA TN-2454, Sept. 1951.

5. Herriot, J. G.: Blockage Corrections for Three-Dimensional-Flow Closed-Throat Wind

Tunnels, with Consideration of the Effect of Compressibility. NACA Report 995, 1950.

6. Braslow, A.L.; and Knox, E.C.: Simplified Method for Determination of Critical

Height of Distributed Roughness Particles for Boundary-Layer Transition at Mach

Numbers From 0. to 5. NACA TN-4363, Sept. 1958.

7. Shevell, R.: and Schaufele, R.: Aerodynamic Design Features of the DC-9. AIAA

Paper 65-738, Nov. 1965.

ll

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TABLE I.- JWRA MODEL GEOMETRIC DATA (Dimensions in feet)

Configuration JW-1 JW-2 JW-3

Span 6.667 5.383 4.000

Length 5.597 5.597 5.597

Height 1.271 1.271 1.271

Sref 3.039 2.651 2.137

Cref 0.484 0.510 0.543

Forward Wing

Area 3.039 2.651 2.137

Sweep(c/4) 30.5 ° 30-5° 30"5°

Taper Ratio 0.4 0.4 0.4

Dihedral 5.0 ° 5.0° 5"0°

Rear Wing

Area 0.933 0.933 0.933

Sweep(c/4) -32.0 ° -32.0 ° -32.0 °

Taper Ratio 0.6 0.6 0.6Dihedral - 20.0 ° - 20.0 ° - 20.0 °

Vertical Tail

Area 0.475 0.475 0.475

Sweep(c/4) 32 ° 32° 32°

Taper Ratio 0.3 0.3 0.3

Datums (fuselage stations)

Nose 0.611 0.611 0.611

Fwd wing L.E. @CL 2.796 2.796 2.796

Rear wing L.E. @C_ 5.661 5.661 5.661

Balance Center 3.986 3.986 3.986

Moment Reference 3.886 3.804 3.695

Main landing gear 4.122 4.122 4.122

Root

Joint

Tip

Root

Joint

Tip

Forward Wing

JWR

JWJ

JWT

7.5 °

5.5 °

2.1 °

Airfoils

Rear Wing

JTR

JTT

Incidence

3.0 °

5.0 °

Vertical Tail

JFR

JFT

0.0 °

0.0 °

12

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TableII.- AIRFOILCOORDINATES

x/c

0.00000.00250.00500.00750.01000.01500.02000.02500.03000.04000.05000.06000.07000.08000.09000.10000.1200O.14000.16000.18000.20000.22000.24000.26000.28000.30000.33000.36000.39000.42000.45000.48000.51000.54000.57000.60000.63000.66000.69000.72000.75000.78000.81000.84000.87000.90000.93000.960O0.98001.0000

JWR (wing root)y/c upper y/c lower

0.00000 0.000000.00953 -0.007760.01350 -0.010370.01641 -0.012280.01883 -0.013830.02286 -0.016150.02624 -0.017960.02924 -0.019510.03196 -0.020890.03679 -0.023360.04 100 -0.025530.04472 -0.027470.04808 -0.029210.05115 -0.030820.05399 -0.032300.05662 -0.033690.06136 -0.036180.06554 -0.038360.06926 -0.040280.07261 -0.041980.07562 -0.043470.07833 -0.044780.08076 -0.045910.08293 -0.046880.08485 -0.047690.08653 -0.048340.08860 -0.049030.09014 -0.049380.09113 -0.049380.09149 -0.048940.09116 -0.047990.09013 -0.046540.08838 -0.044560.08590 -0.042050.08277 -0.039120.07907 -0.035850.07489 -0.032340.07024 -0.028630.06518 -0.024790.05978 -0.020900.05408 -0.017050.04810 -0.013260.04189 -0.00964

0.03552 -0.006280.02904 -0.003300.02251 -0.000850.01602 0.000860.00968 0.001450.00561 0.000860.00145 -0.00145

JWJ (wing join0y/c upper y/c lower

0.00000 0.000000.01191 -0.008340.01695 -0.011020.02063 -0.012990.02363 -0.014580.02851 -0.016980.03250 -0.018890.03593 -0.020530.03896 -0.022000.04416 -0.024630.04850 -0.026930.05221 -0.028960.05544 -0.030790.05831 -0.032460.06088 -0.034000.06321 -0.035420.06726 -0.037960.07071 -0.040150.07368 -0.042060.07629 -0.043740.07857 -0.045210.08059 -0.046490.08236 -0.047590.08390 -0.048550.08523 -0.049350.08635 -0.05001

0.08763 -0.050730.08846 -0.051140.08882 -0.051240.08861 -0.050930.08779 -0.050150.08633 -0.048880.08421 -0.047110.08143 -0.044830.07808 -0.042120.07422 -0.039070.06993 -0.03577

0.06527 -0.032250.06025 -0.028560.05496 -0.024790.04946 -0.021000.04374 -0.017210.03788 -0.013530.03194 -0.01002

0.02598 -0.006810.02007 -0.00402

0.01431 -0.001850.00884 -0.000650.00547 -0.00071

0.00225 -0.00225

JWT (wing tip)y/c upper y/c lower

0.00000 0.000000.01215 -0.008620.01736 -0.011300.02115 -0.013270.02423 -0.014850.02919 -0.017260.03320 -0.019180.03663 -0.020840.03963 -0.022330.04472 -0.025000.04892 -0.027340.05246 -0.029400.05551 -0.031260.05820 -0.032960.06059 -0.034530.06275 -0.035980.06650 -0.038560.06969 -0.040780.07248 -0.042710.07496 -0.044380.07719 -0.045830.07921 -0.047070.08105 -0.048130.08270 -0.049000.08420 -0.049710.08552 -0.050250.08719 -0.050760.08846 -0.050930.08931 -0.050730.08962 -0.050100.08932 -0.048970.08839 -0.047340.08679 -0.045190.08450 -0.04254

0.08159 -0.039470.07815 -0.036080.07422 -0.032480.06985 -0.028700.06507 -0.024820.05993 -0.020930.05449 -0.017110.04875 -0.013410.04277 -0.009920.03659 -0.006760.03027 -0.004030.02387 -0.001910.01750 -0.000590.01129 -0.000410.00736 -0.001260.00360 -0.00360

13

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x/c

0.0000

0.0025

0.0050

0.0075

0.0100

0.01500.0200

0.0250

0.03000.0400

0.0500

0.0600

0.0700

0.0800

0.0900

0.10000.1200

0.1400

0.1600

0.1800

0.2000

0.2200

0.2400

0.260O

0.2800

0.30000.3300

0.3600

0.3900

0.4200

0.4500

0.4800

0.5100

0.5400

0.5700

0.6000

0.6300

0.6600

0.6900

0.7200

0.7500

0.7800

0.8100

0.8400

0.87000.9000

0.9300

0.9600

0.9800

1.0000

Table II.- (CONTINUED)

JTR (tail root)

y/c UPl_ y/c lower

0.00000 0.00000

0.00920 -0.00716

0.01253 -0.01026

0.01499 -0.01257

0.01703 -0.01446

0.02040 -0.017480.02323 -0.01991

0.02575 -0.02204

0.02807 -0.02397

0.03225 -0.02743

0.03597 -0.03048

0.03930 -0.03320

0.04233 -0.03567

0.04512 -0.037920.04770 -0.04000

0.05011 -0.04194

0.05448 -0.04545

0.05834 -0.04851

0.06175 -0.051200.06476 -0.05353

0.06742 -0.05555

0.06974 -0.05728

0.07175 -0.058740.07347 -0.05993

0.07489 -0.06087

0.07604 -0.06157

0.07724 -0.06216

0.07783 -0.06222

0.07781 -0.06176

0.07719 -0.06078

0.07595 -0.05928

0.07412 -0.05727

0.07174 -0.05481

0.06889 -0.05196

0.06561 -0.04878

0.06194 -0.04533

0.05795 -0.04 165

0.05366 -0.03779

0.04912 -0.03379

0.04438 -0.02972

0.03947 -0.02561

0.03447 -0.02154

0.02944 -0.01757

0.02448 -0.01381

0.01967 -0.01035

0.01512 -0.00732

0.01092 -0.00486

0.00721 -0.00320

0.00512 -0.00279

0.00340 -0.00340

(tail tip)

y/c upper y/c lower

0.00000 0.00000

0.00914 -0.00711

0.01251 -0.01022

0.01503 -0.01258

0.01714 -0.01453

0.02067 -0.017700.02365 -0.02030

0.02630 -0.02257

0.02872 -0.02462

0.03309 -0.02829

0.03694 -0.03150

0.04038 -0.03435

0.04350 -0.03691

0.04635 -0.03923

0.04899 -0.041360.05144 -0.04333

0.05586 -0.04685

0.05973 -0.04989

0.06314 -0.05254

0.06613 -0.05480

0.06872 -0.05672

0.07097 -0.05832

0.07289 -0.05961

0.07450 -0.06063

0.07581 -0.061390.07681 -0.06190

0.07775 -0.06219

0.07799 -0.06194

0.07757 -0.06117

0.07654 -0.05988

0.07491 -0.05808

0.07276 -0.05584

0.07012 -0.05322

0.06704 -0.05026

0.06359 -0.04701

0.05981 -0.04350

0.05574 -0.03978

0.05141 -0.03590

0.04689 -0.03193

0.04224 -0.02794

0.03753 -0.02400

0.03281 -0.02019

0.02814 -0.01657

0.02362 -0.01325

0.01933 -0.01031

0.01535 -0.00790

0.01178 -0.00616

0.00873 -0.00527

0.00706 -0.00522

0.00570 -0.00570

14

JFR (fill root)

y/c symmetric

0.00000

0.01285

0.01808

0.02194

0.02501

0.02972

0.03338

0.036480.03919

0.04372

0.04739

0.05044

0.053040.055280.05723

0.05897

0.06194

0.06438

0.066450.06825

0.06982

0.071200.07242

0.07349

0.07441

0.07518

0.07604

0.07655

0.07668

0.07630

0.07534

0.07377

0.07156

0.06871

0.06532

0.06145

0.05722

0.05264

0.04777

0.042710.03750

0.03218

0.02683

0.02153

0.01636

0.01245

0.00845

0.00510

0.00348

0.00250

JFT (f'm tip)

y/c symmetric

0.00000

0.00871

0.01238

0.015110.01729

0.02053

0.02291

0.024900.02664

0.02950

0.03174

0.03353

0.03502

0.036270.03733

0.03824

0.039740.04092

0.04189

0.04271

0.O4341

0.04403

0.04457

0.04504

0.04544

0.04578

0.04614

0.04634

0.04634

0.04609

0.04554

0.04466

0.04342

0.04182

0.03989

0.03766

0.03518

0.03249

0.02961

0.02658

0.02345

0.02022

0.01694

0.01367

0.01050

0.00755

0.00486

0.00260

0.00149

0.00080

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TABLE III.- AVAILABLE CONTROL SURFACE DEFLECTIONS FOR THE JWRA MODEL

Surface Baseline Longitudinal Lateral

right/left

FI 0° 20° -15°/10 °FM 0° 15° -15°/10 °FO 0° -15°/10 °RI 0° 100,-5°,-15 °

RO 0° 100,-5°,-15 o -15°/10 °

RUDDER 0°,15 °

15

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Figure I. NASA AD-I Oblique-Wing Aircraft.

Figure 2. One-Sixth-Scale JWRA Model Installedin the Ames 12-FootPressure Wind Tunnel.

ORIGINAL PAGE

BLACK AND WHITE PHOI'O_

16

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A

(Dimensions are in inches)

45 °

Figure 3. Model Support Strut.

Figure 4. J'W-I Configuration.

17

Page 24: Experimental Aerodynamic Characteristics of a Joined-Wing … · 2014-10-04 · Experimental Aerodynamic Characteristics of a Joined-Wing Research Aircraft Configuration Stephen C

Figure 5. 7#-2 Configuration.

Figure 6. 7#-3 Configuration.

18

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FORO

RI

Figure 7. Control Surface Nomenclature.

19

Page 26: Experimental Aerodynamic Characteristics of a Joined-Wing … · 2014-10-04 · Experimental Aerodynamic Characteristics of a Joined-Wing Research Aircraft Configuration Stephen C

3

(Dimensions are in inches)

375

4.875

Figure B. Vortilon installation.

2O

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(Dimensions are in inches)

J'q_---1.625

I

GE STATION 4.08

65°

Figure 9. Speed Brake Installationon Engine Pylon.

21

Page 28: Experimental Aerodynamic Characteristics of a Joined-Wing … · 2014-10-04 · Experimental Aerodynamic Characteristics of a Joined-Wing Research Aircraft Configuration Stephen C

Zw

NOTES:

x W

CL

1. POSITIVE DIRECTIONS OF FORCE AND MOMENT COEFFICIENTSAND ANGLES ARE SHOWN BY ARROWS.

2. FOR CLARITY ORIGIN OF WIND AXES, xw, Yw, Zw, HAVE BEENDISPLACED FROM THE CENTER OF GRAVITY.

Figure I0.Axis Systems and Sign Conventions for Aerodynamic Data.

22

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N/LRA_alo_ _oroutic5 a_ Report Documentation Page

1. Report No.

NASA TM-101083

4. Title and Subtitle

2. Government Acc_on No.

Experimental Aerodynamic Characteristics of a

Joined-Wing Research Aircraft

7. Author(s)

Stephen C. Smith and Ronald K. Stonum (First Lieutenant,Force Systems Command Liaison Office, U.S. Air Force)

9. Performing Organization Name and Addreu

Ames Research Center

Moffett Field, CA 94035

12. Sponsoring Agency Name and Address

National Aeronautics and Space AdministrationWashington, DC 20546-0001

15. Supplementary Notes

Point of Contact:

3. Recipient's Catalog No.

5. Report Date

April 1989

6, Performing Organization Code

8. Performing Organization Report No.

A-89074

10. Work Unit No.

505-61-21

11. Contract or Grant No.

13. Type of Report and Period Covered

Technical Memorandum

14. Sponsoring Agency Code

Stephen C. Smith, Ames Research Center, MS 227-2, Moffett Field, CA 94035

(415) 694-5856 or FTF 464-5856

16. Abstract

A wind-tunnel test was conducted at Ames Research Center to measure the aerodynamic characteristics of a joined-wingresearch aircraft (JWRA). This aircraft was designed to utilize the fuselage and engines of the existing NASA AD-1 aircraft.The JWR A. was designed to have removable outer wing panels to represent three different configurations with the interwingjoint at different fractions of the wing span. A one-sixth-scale wind-tunnel model of all three configurations of the JWRA wastested in the Ames 12-Foot Pressure Wind Tunnel to measure aerodynamic performance, stability, and congol characteristics.

This report presentsthe resultsof these tests.Longitudinal and latera]-direcdona/characteristicswere measuredoveranangle of attack range of -7 ° to 14° and over an angle of sideslip range of -5 ° to +2.5 ° at a Mach number of 0.35 and aReynolds numberof 2.2 x 106/ft. Various combinationsof deflectedcontrolsurfaceswere tesw.zlto measure the effectivenessand impact on stability of severalcontrol surface arrangements.In addition, the effects on stall and post-stallaerodynamiccharacteristics from small leading-edge devices called vortilons were measured.

The results of these tests indicate that the JWRA had very good aerodynamic performance and acceptable stability andcontrol throughout its flight envelope. The vortilons produced a profound improvement in the stall and post-stall characteris-tics with no measurable effects on cruiseperformance.

17. Key Words (Suggested by Author(st)

Joined wingVortilons

Experimental aerodynamics

19. Security Classif. (of this report)

Unclassified

NASA FORM 1626 OCT 86

18. Distribution Statement

Unclassified - Unlimited

Subject category - 02

I 20. Security Classif. (of this page) 21. No. of Pages ' 22. Price

i Unclassified 98 A051

For sale by the National Technical Information Service. Springfield. Virginia 22161

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