Wind tunnel tests of high-lift systems for advanced transports using high-aspect-ratio supercritical wings

PART I WIDE BODY DC-X-200-TYPE MODEL LB486-C MODEL INSTALLED IN LANGLEY V/STOL TUNNEL

PART I

WIDE BODY DC-X-200-TYPE MODEL

LB-486B,C MODEL DESCRIPTION

The wind tunnel model used for the program was a 4.7-percent representation

of the DC-X-200 aircraft, and was the same as that used in Phase I of the

EET Project study. The model is depicted in Figure 1. The configuration

notation data, dimensional data, and grid position definitions are presented

in Appendixes A, B, and C, respectively. The model was designed as a

primary high-lift configuration that included a variable-camber Krueger

(VCK). Secondary configurations employed either slats or fixed-camber

Kruegers (FCK) along the leading edge.

Combinations of an FCK inboard with

a slat outboard were also tested.

The primary trailing-edge configuration employed inboard and outboard

two-segment flaps. Between these two flaps was a flaperon, essentially a

single-slotted flap, that could be articulated in the same manner as the

DIMENSIONS IN CENTIMETERS (INCHES) MODEL SCALE 36.530 (14.382)

I

70.236 (27.652)

l-

#

222.08 (87.435)

I

FIGURE 1. HIGH-LIFT LOW-SPEED WIND TUNNEL MODEL but that incorporated a high-speed, main flap for the high-lift conditions,

At the

short-chord aileron in the retracted, or cruise, configuration.

this aileron was locked in an undeflected position.

high-lift condition,

This permitted an 83-percent continuous flap span resulting in an improved

span loading for high-lift conditions. The various high-lift components are

depicted in Figure 2.

SECTION A-A SECTION B-B [LEADING EDGE DEVICES/ (TRAILING EDGE DEVICES] PRIMARY CONFIGURATION - VCK PRIMARY CONFIGURATION - TWO-SEGMENT FLAP TAKEOFF AND CLEAN CLEAN LANDING TAKEOFF LANDING \ SECONDARY CONFIGURATION -SLAT SECONDARY CONFIGURATION -SINGLE-SEGMENT FLAP TAKEOFF LANDING SECONDARY CONFIGURATION- FCK FIGURE 2. HIGH-LIFT COMPONENTS EVALUATED IN EXPERIMENTAL TEST PROGRAM

The model also included an aileron on the left wing panel, spoilers, and a

remote-drive horizontal stabilizer deflection capability. Other model

landing gear, and a cruise wing

components included nacelles, pylons,

trailing edge (i.e., flaps retracted). The fuselage consisted of DC-10

model nose and aft fuselage shell sections, and a top center section and

wing/fuselage fillet developed for Phase I testing.

A fuselage core was adapted for attachment of the fuselage shell sections, support of two !&module scanivalve systems, support of a bubble pack plate,

and attachment of the wing and the vertical and horizontal stabilizers. A

fuselage internal pitch system was installed in the core. This system

permited the fuselage to be pitched from aFRP = O" to +lO" while the

internal balance remained at oFRP = O". The other pitch angles were

obtained by using the external pitch system. This system provided more

accurate drag measurements between O" to 10'.

The wing geometry and planform dimensions are shown on the wing diagram

(Figure 3). The wing was designed to simulate the aircraft wing under a l-g

load.

It incorporated the following features:

1. A cruise leading edge removable at the front spar. This leading

edge was tested with and without simulated VCK stowage wells. Also

provided was a WUSS (wing under slat surface) leading edge for the

slat configuration.

2. A VCK, FCK, and slat leading-edge flap device with variable

deflection and position capability.

3. A two-segment trailing-edge flap supported at five deflection

angles by fixed brackets simulating the airplane flap linkage.

Variable position capability was provided for the main flap.

4.

A manually set aileron, left side only, and spoilers both sides.

5.

Approximately 400 static pressure orifices installed in the VCK, slat, wing, and flaps.

The geometry of the horizontal stabilizer is shown in Figure 4. The

horizontal stabilizer was removable for testing tail-off. Each side of the

stabilizer was fabricated in one piece without elevators.

A remote control

system was used to vary the stabilizer incidence between +5O and -15'.

The vertical stabilizer planform is shown in Figure 5. The stabilizer was

fabricated as one piece without rudders and was removable to provide a

tail-off configuration.

- 121.964 146.025l 11,677 7) f (4.5971 1: I%:..,,,, / = 36.367 xv4 DIMENSIONS IN CENTIMETERS (INCHES) MODEL SCALE 114.318) 28.744 I / / Ill.3171 / lC/4)MAC I Y = 160.280 163.1021 .I -. Fl.. ,.^ ^_A, 7 I 37.761 14.8671 FRONT SPAR PLANE 51.895 120.431) /...-- -1 .i ” -.-- I\ 6.954 12.7381 8.692 13.4221

~““I~“~“Y3r”ILtnb

L---- 30.793 (12.1231 ---__--- 43.411 I1 7.091)

I-

OUTaOARO TWO.SEGMENT FLAP REAR SPAR PLANE ‘- LOW-SPEED AILERON FIGURE 3. WING (w,,) DIAGRAM = 0.1298 ITI* 11.397 FT*) -Y = 231.87 (91.287) SH Ffl = 3.80 x = 0.350 SWEEPCH = 30’ r = lo.o” 0.781) DIMENSIONS IN CENTIMETERS (INCHES) PlVflT AYIC KR I\6

...U, -,.I” I”.“”

MODEL SCALE -PERCENT CR 4.536~J .7K6) ABOVE FRP YH = 16.11 (6.344) 9.596 ($.778)1 \MODEL ACTUAL TRAILING EDGE (CUT BACK TO ACHIEVE 0.03 (0.010) THICK TRAILING EDGE) FIGURE 4. HORIZONTAL STABILIZER (HIA) DIAGRAM 12.87 (5.065) Sv = 0.09850 III* (1.0663 FT*) t- pi = 1.600 h = 0.35 SWEEP Cv = 35O DIMENSIONS IN CENTIMETERS (INCHES) MODEL SCALE MACV = 26.731 (10.524) THEORETICAL TRAILING EDGE ACTUAL MODEL TRAILING EDGE (CUT BACK TO ACHIEVE 0.03 (0.010) THICK TRAILING EDGE) 36.759 (14.472) FIGURE 5. VERTtCAL STABILIZER (V,A) DIAGRAM

Flow-through nacelles (Figure 6) from a DC-10 model were used and were

The pylon plane of symmetry had a 1.8O

attached to the wing by pylons.

toe-in relative to the airplane plane of symnietry (measured in the FRP) and

was perpendicular to the FRP with the wing in a rigged position with a

dihedral angle of 4.05'. Nacelle strakes were attached to the nacelle for

most tests.

DIMENSIONS IN CENTIMETERS (INCHES) MODEL SCALE 4.30 (1.69)

-I I-

3.25 PERCENT CHORO

I-b-

ENGINE CENTERLINE AT+1.6’lNClDENCE TO THE FRP LEXISTING DC-10 GE NACELLE (N2A) FIGURE 6. NACELLE/PYLON (Na Pm) DIAGRAM

The nose gear simulated the DC-10 nose gear in structure and location. The

main landing gear simulates the airplane gear configuration with oleos

Extended main gear wheel well cavities were not simulated. A

extended.

retracted main landing gear configuration was also provided.

The definitions of gap, overhang (O.H.), and deflection used to position the

leading-edge high-lift devices are illustrated in Figure 7. The deflection

angles were measured in a streamwise plane oriented normal to the wing

reference plane (WRP). Definitions for main and aft flap gap, O.H., and

deflections are shown in Figure 8. The same definitions were used for both

The variable test positions tested are

the flaperon and the main flap.

defined and identified in the grid notations table of Appendix C.

CLEAN WING MAX LENGTi-l LINE VCK. FCK MLL / bRACKET FIGURE 7. LB-486 LEADING EDGE DEVICE GAP, OVERHANG, AND DEFLECTION DEFINITIONS v FIGURE 8. LB-486 FLAP GAP, OVERHANG, AND DEFLECTION DEFINITIONS FIGURE 8. LB-486 FLAP GAP, OVERHANG, AND DEFLECTION DEFINITIONS

LB-486B,C INSTRUMENTATION

Aerodynamic forces on the model were measured using the Ames Task Mark II

10.16-cm (4-in.) diameter internal balance at the Ames l&Foot Pressure Wind

For the NASA Langley V/STOL Wind Tunnel

Tunnel (LB-486B test).

(LB-486C test), the balance used was the Langley 5.08-cm (Z-in.) diameter

internal balance.

In the Ames test, electrolytic alignment bubbles housed in the fuselage nose

were used to measure the angle of attack of the fuselage reference plane.

From angles of attack of -6O to O", the model was pitched by the

From O" to +lO" angles of attack, the fuselage

external pitch drive.

was pitched using the fuselage internal pitch drive while maintaining the

balance at 0'. For angles of attack of 10' to 34O, the fuselage was

pitched using the external pitch drive with a loo angle maintained between

the balance axis and the fuselage axis.

In the Ames test the horizontal stabilizer incorporated remote drive and

dual-position potentiometer for changing tail incidence during a run. In

the NASA V/STOL test, a NASA-furnished electronic inclinometer was used to

determine angle of attack. The horizontal-tail incidence in the V/STOL test

was set at O".

LB-486B,C MODEL INSTALLATION

The model was installed in the NASA Ames 12-Foot Pressure Wind Tunnel on the

The model was pivoted about the

tandem support system shown in Figure 9.

main strut pivot point and was powered by the aft pitch strut. The entire

air loads on the strut are not sensed by

strut system was nonmetric (i.e.,

the balance). The struts entered the fuselage as far aft as practical to

minimize the aerodynamic interference effects on the model.

TS BALANCE 306.616 (120.715) CENTER I -- - - -- s 2.985 - 29.57 (1.175) (11.64) AMES TASK MK II BALANCE DATAREFERENCE CENTER oiM~r~s~0r4s IN CENTIMETERS (INCHES) MODEL SCALE FIGURE 9. MODEL ItiSTALLATlON IN THE NASA AMES 12-FOOT PRESSURE WIND TUNNEL

The same support system (Figure 10) was utilized during the NASA Langley

V/STOL test program. It was adapted to the existing V/STOL Tunnel

structure; extensions for the main and pitch struts were added to the basic

tandem strut system. The extensions permitted the model to be located near

the vertical position of the tunnel centerline.

REVIEW OF PHASE I RESULTS

During Phase I, the aerodynamic characteristics of the clean wing, VCK,

slat, and flaps were defined experimentally. The lift and pitching-moment

curves for the clean wing are shown in Figure 11. These curves indicate

that the cruise wing, as defined for Phase I, was subject to outboard stall,

although it is likely that the curves overstate the tendency for stall

because of the Reynolds number effect. Because of the short tip chord of

the wind tunnel model, the highest Reynolds number condition resulted in a

tip chord Reynolds number of only 1.9 million. Figure 12 shows that higher

TS = 506.43 (200.17) MS = 160.28 (63.102) MS = 65.301 (25.709) I LANGLEY 748 BALANCE ,i , ~P~TCHSTRUT BALANCE CENTER- /m I 8AYONET- MAIN STRUT

I

EXTENSION PITCH STRUT c= EXTENSION -I- I M ($ MAST SUPPORT TOP SURFACE

I

V/STOL TUNNEL FLOOR- PITCH DRIVE MECHANISM I DIMENSIONS IN CENTIMETERS (INCHES) MODEL SCALE FIGURE 10. MODEL INSTALLATION IN THE NASA LANGLEY V/STOL WIND TUNNEL

's for the outboard wing

stall angles and larger values of section C

LMAX

panel might have been obtained if the test could have been made at a higher

Later high-aspect-ratio supercritical wing designs have

Reynolds number.

shown improvements in stall angles and C

LMAX'

l .

i SYM RUN MODEL LB466A Ii f3 ” a CONFIGURATION S-N-- P-m te _ V. _ H. _ 85 1 OFF t 0 1 23 i 1 1 ZA ZA 1A 1A 1A ;i 0. YOO p& MACH = 020 3. 50- RNMAC = 5.12 x lo6 I 3. oo-; I I I -10 30 I- 2. 00 I. 50 -0.200 I3 ANELE OF RTTRCK-DEG I. 00 -0.300 -0. L100 0. 50 0” - I I I I I I -5; 5 IO I5 20 25 30 0” B RN6LE OF RTTRCK-DEG / -0. 50 .~- LIFT AND PITCHING MOMENT FIGURE 11. TAIL-ON CHARACTERISTICS FOR THE CRUISE WING WITH NACELLES, PYLONS, AND STRAKES ATTACHED MODEL LB-488A CONFIGURATION S, NzA PPAZIA MACH = 0.20 1.

1 .c 1 5 10 20 RN x 10 MAC FIGURE 12. EFFECT OF REYNOLDS NUMBER ON CLEAN-WING SECTION MAXIMUM LIFT

Figures 13 and 14 show the lift and pitching-moment characteristics for the

primary VCK and slat configurations tested. While the C and L/D ratio

LMAX

for the slat configurations were marginally better than those of the VCK

configurations, use of the VCK resulted in superior stall characteristics.

Configurations including slats exhibited both pre-stall and post-stall

nose-up tendencies. While the VCK configurations showed post-stall nose-up

trends, the pre-stall characteristics were good. Nearly all of the work

accomplished on this model during Phase II was directed toward improving the

low-speed stall characteristics by making adjustments in leading-edge device

position and type.

The trailing-edge flap studies of Phase I indicated that the changes in

performance due to gap and overhang variations were not as significant as

the corresponding variations for the leading-edge devices. As expected, the

CLMAX and

two-segment flap was superior to the single-segment flap in

flap

lift increments. Trimmed polar comparisons indicated that the

single-segment and two-segment flaps resulted in equivalent L/D envelopes

for takeoff flap settings. For equivalent values of approach speed, the L/D

values for the two-segment flap were superior to those of the single-segment

flap.

Because of these definitive results, little additional flap

In

optimization work was conducted on the wide-body model during Phase II.

addition to the high-lift work, Phase I testing also defined the

effectiveness of the spoilers and ailerons.

MODEL LB466A i RUN H CONFIGURATION S4HIAVIA MACH = 0.20 OFF 116 O0 121 3. 50.

RNI\IIAC = 5.12 x lo6 -5O 122 6 VCK = 45El45G = 5CllOB 3. 00 ‘FLAP 2. 50 I Qc@3 I I I 1 I I -5 5 20 25 30 RNGLQO:‘flTTRCK-D;; 2. 00 -0. 100 a I. 50 -0. 200 B 0 o Q I. 00 0 -0. 300 -0. qoc C&O -0.5oc - ; 1’0 1’5 ;0- 25 -0. hoc RNGLE OF RTTRCK-DEG -0. 7oc LIFT AND PITCHING MOMENT FIGURE 13. TAIL-ON AND TAIL-OFF AERODYNAMIC CHARACTERISTICS OF THE VCK WITH TWO-SEGMENT TAKEOFF FLAPS CONFIGURATION MODEL LB-466A BASIC CONFIGURATION S4HIAVIA MACH - 0.20 3.50-, RN = 5.12 x 10’ MAC 6 - 15D125D SLAT Y 6 = 5CllOB 2 I: 0.200- 3. 00 FLAP i I 2.50-l jo

:J;r -o.boo

-0. 700 i -._.-- LIFT AND PITCHING MOMENT FIGURE 14. TAIL-ON AND TAIL-OFF AERODYNAMIC CHARACTERISTICS OF THE SLAT WITH TWO-SEGMENT TAKEOFF FLAPS CONFlGURAtltJN c:

LB-486B,C RESULTS AND DISCUSSIONS

Most of the work on the wide body model during Phase II was directed toward

improving the pitching-moment characteristics of the wing, without causing

The approach consisted of either increasing

an excessive loss in C

LMAX'

the stalling angle of the outboard wing panel, or tuning the stall angle of

the inboard wing to be just below that of the outboard wing. Additionally,

to prevent post-stall pitch-up, it was desirable that the stall inboard be

due to separation at the leading edge of the high-lift device, thereby

increasing the rate of lift loss inboard relative to that outboard.

Configurations tested included a VCK with a reduced deflection, trimmed

slats inboard, a normal-chord and a short-chord FCK, a differential flap

deflection, and a two-segment flaperon. In addition to the study of these

configurations designed to improve C and/or pitching-moment trends,

LMAX

the improvement in takeoff L/D performance due to sealed slats was

evaluated, the penalty associated with use of a high-speed aileron was

determined, and data obtained at the Langley and Ames tunnels were compared.

Reduced VCK Deflection

Phase I results (LB-486A) showed equivalent C values for the slat and

LMAX

VCK configurations. However, the lower minimum pressure coefficients on the

VCK indicated that a reduction in deflection might delay leading-edge

separation and result in increased maximum lift. A VCK deflection of

= 33' compared to the Phase I value of 6VCK = 45" was

&VCK

therefore selected for the LB-486C test at the NASA Langley V/STOL

Facility. Results of this test indicated that it was not possible to obtain

increased C due to the low Reynolds number (1.14 million) available in

LMAX

this tunnel. Further examination of the configuration was made at a higher

Reynolds number (5.89 million) during the Ames 12-Foot Tunnel entry

(LB-486B).

The same results as in LB-486C were observed. The reduced

deflection resulted in a lower outboard stall angle than the 45O

The basic 45', 33', and 45O/33O (inboard/outboard) VCK

deflection.

deflection lift and pitching-moment data are shown in Figure 15. The

corresponding drag values indicated L/D values at 1.3Vs of 9.52, 10.0, and

and 45O/33O VCK deflections, respectively.

9.0 for the 45', 33', MODEL LB4866 ;YM 1 RUN t Z CONFIGURATION B 2A w3B ‘2B N2A ‘2A 1A MACH = 0.20 RN = 5.12x lo6 3. 50. MAC 6 = 25Kl12C FLAP 3. 00.

B

(P

2. 50.

P

0 0 RNGLE OF RTTRCK-DEG 2. 00.

r3 q -0.200 -I I. 50.

-0. 300

-0. YOO 1

I. 00.

0. 50.

3 I I I I I I 5 IO 15 20 25 -5 * RNGLE OF RTTRCK-DEE -0. 50.

NSTC A. LIFT AND PITCHING MOMENT FIGURE 15. EFFECT OF VCK DEFLECTION -9 9. 0 8. 0 I- 7. 0 x .

i : I- $ b.0 L b : I- 3 5.0 .Y ,- 4. c I- 3. c ,- 2. c ,- I. c b-f- --ed I I I I I ! I I I I 1 I I 1 I I 1 C. 04 t 3. ( 10 0. 04 0. ox 0. I.2 0. lb 0. 20 0. 24 0. 2x 0. 32 0. 3b 0. 40 0. 44 0. 4s 0. 52 0. 5b 0. b0 ‘0. b4 0. bE 72 0.

DRRG COEFFICIENT NSTC 0. DRAG FIGURE 15. EFFECT OF VCK DEFLECTION

Sealed Slats

In the Phase I LB-486A tests, a landing slats/takeoff flaps combination was

investigated since it would simplify the high-lift system mechanically to

have only one slat position for both takeoff and landing. The results

showed, however, that the landing slat reduced L/D when used with either a

clean trailing edge (GFLAp = 0') or the basic takeoff flap deflection

To improve the L/D for this combination a sealed

= 5O/lOO).

(&FLAP

(i.e., zero gap) inboard and outboard slat configuration was investigated.

The configuration was tested first with a 5O slat deflection inboard and a

Then because previous analysis had shown a

20' deflection outboard.

retracted slat might improve the pitching-moment characteristics, it was

also tested with a 0' deflection inboard and a ZOO deflection outboard.

The results are presented in Figure 16. Because the loads on the sealed

slat were expected to be high, it was not tested at the high Reynolds

number. The results indicate that, as expected, the 50/20°

These were improved

configuration had adverse pitch-moment characteristics.

by retracting the inboard slat, without reducing C

LMAX'

Also shown in Figure 16 is the landing slat configuraton with takeoff

associated with the sealed slat is obvious.

flaps. The CLMAX penalty

Figure 16 shows the O"/200 slat configuration gave slightly higher L/D

than the 50/20° slat configuration, tail-on. Tail-off L/D's for clean,

sealed, and slotted configurations are compared in Figure 17. The improved

tail-off L/D values for the sealed configuration at 50/10° flap

deflection are illustrated. High Reynolds number data for the clean

trailing edge with sealed slat configuration were not obtained.

An inboard sealed slat deflection of 5' was tested with landing flaps and

an outboard landing slat position.

Results indicated a substantial C

LMAX

degradation and post-stall nose-down pitching-moment trends (Figure 18).

LB-486A testing included a 15' inboard sealed slat position; the results

showed no adverse effects on C

and no change in pitching-moment

LMAX

characteristics. An inboard sealed or small-gap slat configuration at an

intermediate inboard slat deflection is a candidate for future low-speed

studies.

“. I..” I- MODEL LB486B 6 RN 6 SYM RUN SLAT MAC CONFIGURATION I3 =I 2A w3EI NPA ‘ZA’IA H1A “IA H 0. 300- k w 2.89 x lo6 5 SEALED120 SEALED 0 42 MACH = 0.20 3. 50 s lo6 cl FLAP F 1 6 = 5CllO 5 8 2.89 5.11 x x 0.200- lo6 CLEAN/SO 15 SLOTTED125 SEALED SLOTTED V 46 41 V -0. 00 $1 RNGLE OF RTTRCK-DEE -0. 200 @cl B vv v v v v i

-0. 500 1 0

q q I7 -0. boo i

I3

q -0. 700 RNGLE OF HTTRCK-DEG -0. 50 -0. 800 J NSTC A. LIFT AND PITCHING MOMENT FIGURE 16. EFFECT OF SEALED SLATS 9. o-

MODEL LB-‘tXb B

RN MAC 6 SYM RUN SLAT v v v 8. o- .

V V 2.89 x lo6 5 SEALED/PO SEALED 0 42 v 2.89 x lo6 CLEAN/PO SEALED Cl 46 v 7. o- ; 5.11 x lo6 15 SLOTTED125 SLOTTED V 41 v P

q @ El

q 0

$3

4. 0

w

q Q El q q q 3. 0 2. 0 I I I I I 1 I , , I 1 1

0. 2Lt 0. 28 0. 32 0. 3b 0. 40 0. 44 0. 4x 0. 52 0. 5b 0. b0 : 0. b’t 0. b8 ( 1. ; 72

DRRE COEFFICIENT NSTI:

I

B. DRAG FIGURE 16. EFFECT OF SEALED SLATS MODEL LB486 TEST 6 6 FLAP SYM RUN CONFIGURATION B SLAT 2A w3 ‘2B N2A ‘2A ‘IA MACH = 0.20 LB-486A CLEAN/CLEAN 010 D 24 l8- RN LB-486A 15 SLOTTED/25 SLOTTED 5/10 0 178 = 2.58 x lo6 MAC LB-486B CLEAN120 SEALED 5110 v 51 lb- El El v El q V -B v q v a v B q 0 v 12- v e v 0 J io- &?

v 0 v 8- v El q v b- v F v v v v I I I I I I I 1 I I I I I I I I ‘%O 0.2 O.L( 0.b 0.8 I.0 1.2 I.L1 I.b I.8 2.0 2:2 2. L1 2. b 2. 8 3. 0 3. 2 3. q b -21 LIFT COEFFICIENT FIGURE 17. L/D COMPARISONS FOR CLEAN WING, SEALED SLAT, AND LANDING SLAT CONFIGURATIONS MODEL LBQB6B CONFlGURATliIN 6 SYM RUN 2A w3B ‘2B N2A ‘2A ‘IA H,A “IA SLAT :: 0.300- MACH = 0.20 :: 15D125D 23 ki 0 = 1.14x lo6 ” RNhlAC 5 SEALEDl25D 0 24 6 = 25112 2 0.200- FLAP 2 :” I: 0000 J 0. IOO- 0 E 0 QOOO =I a io 1’5 i0 25 j0 RNGLE OF RTTACK-DEG El q -0.200- -0. boo- I I I I , 5 IO I5 20 25 -0. 700- RNGLE OF RTTRCK-DEG -0, 50

I

-0.800 I3 El NSTC A. LIFT AND PITCHING MOMENT FIGURE 18. EFFECT OF INBOARD SEALED SLAT WITH LANDING FLAPS 0 0 MODEL LB-486B 8. 0. 0 7. 0

B

x El * B ; El : q El I: b. 0, L 5 2 5.0 NSTC DRRG COEFFICIENT B. DRAG FIGURE 18. EFFECT OF INBOARD SEALED SLAT WITH LANDING FLAPS

Fixed-Camber Krueger

A fixed-camber Krueger (FCK) is an attractive high-lift device option,

especially inboard, because of its mechanical simplicity and the need to

stall the inboard wing panel just before the outboard panel stalls. The

capability of a very efficient slat or VCK is not needed. As shown in

Figure 19, the full-span FCK produced lift and pitching-moment

characteristics equivalent to those of the full-span slat and full-span VCK

configuration. Use of an FCK inboard with a slat outboard, however,

resulted in improved pitch characteristics (Figure 20). Even though the

FCK/slat combination caused pitch-up to start at a lower angle of attack

than the FCK/FCK combination, pre-stall nose-up tendencies were greatly

reduced, and could possibly be eliminated with additional tuning.

Post-stall characteristics continued to be unsatisatifactory, indicating a

lack of leading-edge separation on the FCK.

To further improve pitching-moment characteristics, a short-chord FCK was

fabricated and tested during the LB-486B series. The chord ratio for this

device was 0.068, extrapolated to the side of the fuselage, and 0.105 at the

leading-edge break (pylon position). The comparable values for the slat

were 0.1803 and 0.1295, respectively. The bulb shape was tailored such that

an inboard, leading-edge stall would be obtained. FCK deflections of 50'

and 70' were evaluated with zero gap and overhang. Examination of the

trailing-edge pressures indicated that a premature inboard stall was being

obtained. Favorable pitch characteristics at stall were obtained

(Figure 211, but at the expense of a substantial reduction in C values

LMAX

of -0.457 and -0.412, respectively, for the two FCK deflections. Shims were

fabricated at the tunnel to obtain a small gap and negative overhang for

this leading-edge device.

The best FCK/slat configuration resulted in

higher maximum lift values and better pitching-moment trends then did the

full-span slat configuration (Figure 22).

Tail-off drag values indicated

L/II values at 1.3Vs of 9.71 and 9.77 for the FCK and basic slat

configuration, respectively.

MODEL LB-499C CONFIGURATION B 2A w3Fl ‘ZB N2A ‘2A ‘IA MACH = 0.20 RN = 1.14x lo6 MAC .3 6 = 25112 FLAP .2 l- -0 ANGLE OF ATTACK (DEG) -.

-0 I 1 1 1 I I I 4 0 Lf 8 12 16 20 2L1 28 ANGLE OF ATTACK (DEG) as It A. LIFT AND PITCHING MOMENT FIGURE 19. EFFECT OF LEADING EDGE DEVICE WITH LANDING FLAPS 18- MODEL LB-486C 16 - 111- 12 - 10 - LID 8- 8- t I I I I I I I I I t I I I I I 2.8 310 24 2.8 1.8 2.,-J 2.2 1.q 1.6 OO -8 1.0 1.2 .2 .It .8 LIFT COEFFICIENT B. LIFT-DRAG RATIOS P I FIGURE 19. EFFECT OF LEADING EDGE DEVICE WITH LANDING FLAPS

R MODEL LB-466C

CONFIGURATION B 2A w3B ‘2B NZA ‘2A ‘,A ‘IA H1A MACH = 0.20 m .> = 1.14 x lo6 RN MAC 3.2 = 25112 FLAP .2 2.8 I I I I I I I q 6 12 16 20 2Y i8 ANGLE OF ATTACK - DEG I I I I I I I -0 I -.q 0 Lt 8 12 16 20 24 28 -.6 ANGLE OF ATTACK-DEG -.Lf -.7 A. LIFT AND PITCHING MOMENT FIGURE 20. COMPARISON OF FULL-SPAN FCK AND FCK/SLAT COMBINATION MODEL LB486C 18 - 16- 14 -

12-

L/D

10 - 2- I I I I I I , I I I I I I I I 2.6 2.8 3.0 OO 1.6 2.0 2.2 2.q 1-2 I.'! 1.6 .q .6 -8 1-o .2 LIFT COEFFICIENT B. LIFT-DRAG RATIO FIGURE 20. COMPARISON OF FULL-SPAN FCK AND FCK/SLAT COMBINATION MODEL LB-466B :: CONFIGURATION B 2A w36 ‘PI3 N2A ‘2A ‘IA ‘lAHIA Ii 0.300 k MACH = 0.20 :: u RN = 6.1 x IO6 3. 50 MAC 3. 00 2. 50 RNGLE OF RTTRCK-DEG 2. 00 -0. 200 I. 50 1 -0. i oo- I. 00 -0. +00-Q Q @Q OR -0. 500- 0. 50 I I I 1 1 I I -0. 700 -5 5 IO 15 20 25 i RNELE OF RTTRCK-DEE

J

-0. 50 -0.800 NSTC A. LIFT AND PITCHING FIGURE 21. GRID STUDY FOR SHORT-CHORD FCK/SLAT COMBINATION 7. o- 0 MODEL LBQBBB x. o-’ Y. o- I I I I I I I 1 I 1 I I I I I 1 I I I OLt 0. OS 0. ox 0. 12 0. lb 0. 20 0. 24 0. 28 0. 32 0. 3b 0. YO 0. YY 0. $X 0. 52 0. 5b 0. b0 0. bY 0. bZ 0 72 * 0.00 DRRG COEFFICIENT NSTC

e

B. DRAG FIGURE 21. GRID STUDY FOR SHORT-CHORD FCK/SLAT COMBINATION MODEL LB-4BBB v-66 CONFIGURATION B 2A w3B ‘2s N2A ‘2A ‘IA “IA HIA t :: MACH = 0.20 =I 0. 300 = 5.11 x lo6 k RN MAC 3. 50 _ 3. oo- 7- ; IO ;5 20 25 30 2.50- a RNGLE OF RTTRCK-DEG 0 0 -0. 20( -0. 3oc RNGLE OF P.TTRCK-DEG -0. 5c NSTC A. LIFT AND PITCHING MOMENT FIGURE 22. COMPARISON OF FULL-SPAN SLAT AND SHORT-CHORD FCK/SLAT COMBINATION MODEL LB486B GI 6.0 q FCKISLAT 70Di26D 0 35 3.0 - 2.0 - 1.o I.

- -.

1 , 1 1 --- 1 \ I I 1 I 1 1 , L

-0.e 0.52

0.04 0.08 0.12 0.16 0.20 0.24 0.28 0.32 0.36 0.40 0.44 0.48 0.56 0.60 0.64 0.68 NSTC DRAG COEFFICIENT B. DRAG FIGURE 22. COMPARISON OF FULL-SPAN SLAT AND SHORT-CHORD FCK/SLAT COMBINATION

Slat Trim Effects

The lift and pitching-moment characteristics for the revised slat trim are

presented in Figure 23. The basic trim consisted of a side-of-fuselage

inboard trim and a sealed over-the-pylon configuration (i.e., continuous

,over the pylon). This base case resulted in a C value of 3.2.

LMAX

Figure 23 also illustrates two other trim variations which showed a C

LMAX

reduction of approximately 0.20. For the first variation, the slat trim was

moved outboard 2.25 cm (1 in.) from the fuselage side. This resulted in

improved pitch characteristics at the stall angle, but pitch-up at

post-stall conditions. In the second variation, in addition to the revised

inboard slat trim an over-the-pylon island (i.e., undeflected slat) trim was

tested. Pitching-moment characteristics similar to those of the basic trim

resulted but with reduced magnitude of pitch-up. Small effects were noted

on L/D performance for the two slat-trim revisions. Examination of

Figures 22 and 23 indicates a lower C and more adverse post-stall

LMAX

behavior for the slat trim configuration than the short-chord FCK.

High-Speed Aileron

In order to determine the benefit of a flaperon, a configuration using a

high-speed aileron in place of the flaperon was tested at the maximum

The results indicated a reduction

landing flap deflection of 35O/12'.

of 0.315 in CLa = The drag increase at 1.3Vs

o and 0.216 in C

LMAX*

was 0.008. High-angle-of-attack pitch characteristics were essentially

similar to those of the basic configuration.

Two-Segment Flaperon Replacement

For several runs, the single-segment flaperon was replaced with a

two-segment flaperon. The effects of the change were evaluated at landing

and takeoff flap deflections. The increases in corresponding C values

LMAX

were 0.061 and 0.039, respectively. Small changes in pitching moment were

also indicated. The drag values indicated essentially no change due to the

two-segment replacement for the single-slot flaperon.

I /SIDE OF FUSELAGE TRIM (BASIC) MODEL LB-4666 c z CONFlGURATlOti B OUTBOARD TRIM 2A w313 ‘2B N2A ‘2A ‘IA HI A “IA =1 l-4 0.300- MACH = 0.20 : CONTINUOUS OVER = 6.11 x lo6 i?i PYLON (BASIC) 3. 50- RNMAC 2 0. 200- 6 = 26112 r;!

FLAP ?

3. oo- x 0. IOO- z b RNGLE OF RTTRCK-DEG 2. oo- w I z I =1 % e ki 1.50- 8 0 8-l t t: El BASIC (SIDE OF FUS AND OVER PYLON) oooooo -0. 500- ITI oooooD~ -0. bOO- 0 . I I I 1 I 1 -5 cr 5 IO I5 20 25 -0. 700 RNELE OF RTTRCK-DEG I -0.50- -0. zooI NSTC -- -- A. LIFT AND PITCHING MOMENT FIGURE 23. EFFECT OF SLAT TRIM WITH LANDING FLAPS 9. 0, x. 0 Y. 0 SLAT TRIM CONFIGURATION SYM RUN 3. 0 2. 0 I. 0 @ I I I I I I I I I I 1 I I I I I I I 09 . 0.00 0. OLt 0. OX 0. 12 0. lb 0. 20 0. 2Y 0. 2x 0. 32 0. 3b 0. Lto 0. w 0. YX 0. 52 0. 5b 0. ho 0. bY 0. bX c 72 DRRG COEFFICIENT NSTC B. DRAG FIGURE 23. EFFECT OF SLAT TRIM WITH LANDING FLAPS

Differential Flap Deflection

A 35'/12' (main flap/auxiliary flap) inboard flap deflection combined

with a 25'/12O outboard flap deflection was also tested to determine the

effect on the low-speed characteristics. Results compared with those of the

basic 25O/12' two-segment flap deflection indicated a small reduction in

CLMAX (-0.046) and slightly more positive pitching moments. The increased

inboard flap deflection did not produce a smaller inboard stall angle and

the associated stall improvements. The differential flap deflection did

result in a drag increase of 0.0180 for the C, range of interest.

Ames 12-Foot and Langley V/STOL Tunnel Comparisons

During the Phase I wind-tunnel tests in the Ames l2-Foot Pressure Tunnel,

several configurations were tested at high Reynolds number as well as at

atmospheric conditions. Two of these configurations were also tested in the

Langley V/STOL facility for comparison. The tandem strut support system was

utilized in both cases. Figure 24 presents the lift and pitching-moment

comparison at the atmospheric condition for the slat with two-segment

takeoff flap configuration. The data presented have been corrected for

tunnel wall effects, but not for strut tare effects since these would be the

same for both wind tunnels. Good agreement between the Ames and Langley

data is shown for the lift coefficient up to the angle of attack for stall.

Sane differences are noted in the post-stall region. The pitching-moment

data show differences for most of the angle-of-attack range. This was also

typical of the VCK configuration used for comparison. Comparison of the

drag characteristics indicated differences of 0.0050 to 0.0070 for the

configurations evaluated. The Ames wall corrections are considered a

possible source of these differences.

MODEL LB4B6B & C MACH = 0.20 = 1.14x lo6 RN MAC 3.2 1 HORIZONTAL TAIL-OFF LANDING GEAR OFF 2.8 2.4 RNGLE OF RTT 1.2 SLAT (15’/25’) +TWO-SEGMENT FLAP (5°/100) 0.8 , * .4.

-0.5 -

p6

-FY- d 4 6 li 16 io i4 & -0.6 - RNGLE OF RTTRCK-DEG I -0.4 LIFT AND PITCHING MOMENT FIGURE 24. AMES 12-FOOT AND LANGLEY V/STOL COMPARISON

PART II NARROW BODY ATMR-TYPE MODEL L-B-507A.M’ObEL INSTALLED IN AMES i2-FOOT PRESSURE TUNNEL

PART II

NARROW BODY ATMR-TYPE MODEL

LB 507A MODEL DESCRIPTION

A 5.59-percent-scale full-span model of the ATMR aircraft was used for this

i

program. This model is shown in Figure 25. The configuration notation

data, dimensional data, and grid position definition are presented in

Appendixes D, E, and F, respectively. The model included a

high-aspect-ratio supercritical wing, variable-position leading-edge slats,

an inboard short-chord FCK, two-segment trailing-edge flaps, wing and

high-lift surface pressure instrumentation, and a remotely driven horizontal

stabilizer. The outboard ailerons and wing spoilers also had deflection

capabilities. The model instrumentation was equipped with the Douglas

internal pitch system. This system was used in conjunction with the Douglas

tandem support system and the Task MK IIC internal strain-gage balance.

The model fuselage utilized the LB-506A (high-speed EET model) nose section

and glass fiber wing/body fillet. These parts were combined with a new

aluminum centerbody and aft section. The constant-diameter hollow center

section was machined on the upper and lower surfaces and internally to

provide clearance for the Douglas 10.16-cm (4-in.) balance housing and

internal pitch system. Other instrumentation housed in the fuselage

included two 6-pat scanivalve modules in the nose, two electrolytic bubbles

measuring the angle of the balance axis, and an electrolytic bubble pack to

measure the fuselage'angle of attack.

The wing for this model (Figure 26) consisted of right- and left-hand panels

which were joined together and to the fuselage by means of a wing splice

plate. The wing had removable leading and trailing edges to allow for the

attachment of high-lift devices, and had movable control surfaces. The wing

also included pressure instrumentation at four spanwise locations, and had a

trailing-edge pressure port at one inboard span location. A diagram of the

high-lift system and the lateral control surfaces is provided in Figure 27.

MODEL LB-507A DIMENSIONS IN CENTIMETERS (INCHES) MODEL SCALE UNLESS OTHERWISE SPECIFIED X MAC = TRAPEZOIDAL AREAS: S = 0.464 III’ (5.000 FT*) REF = 0.113 rn* (1.231 FT*) sli = 0.086 III* (0.931 FT*)

I

% 136.976 Y- = (53.888) - c/4

t-

100.952 - -L = V (39.745) I 13.485 (5.309) FIGURE 25. LB-507A MODEL THREE VIEW MODEL LB-507A TRAPEZOIDAL WING CHARACTERISTICS S = 0.464 m* (5.000 FT*) REF A = 26.OOODEG c/4 T.R.

= h = 0.275 #I = 11.10 MAC = 2.103 cm (8.922 IN.)

b/2 = 113.532 cm (44.698 IN.)

DIHEDRAL = I- = 5.000DEG DIMENSIONS IN CENTIMETERS (INCHES) MODEL SCALE UNLESS OTHERWISE NOTED

l

12.670 (4.988) L FIGURE 26. LB-507A WING IW,,) MODEL LB-507A DIMENSIONS IN CENTIMETERS (INCHES) MODEL SCALE 1 L INBDSLAT \ UI!l OUTBD SLAT L,,

r

OUTED SPOILER fpA INBD SPOILER 3 L X f 1A FIGURE 27. HIGH-LIFT AND LATERAL CONTROL SURFACES

fillet which was developed for the high-speed

'The model utilized the XIB

The glass fiber fillet was modified on the lower surface to

Model LB-506A.

provide access holes for the Douglas tandem support system.

The model was equipped with one set of inboard and one set of outboard

The slats were attached by rigged brackets to a WUSS

leading-edge slats.

leading edge which was interchangeable with the cruise leading edge.

Brackets were available to rig the inboard slats at three different

positions. At one of these three positions, a set of shims could be

installed between the slat brackets and the wing to provide a fourth slat

grid position. The definitions of slat gap and overhang are shown in

Figure 28 (which is Figure 7 repeated for convenience), the various slat

SLAT OVERHANG (-I SHOWN CLEAN WING MAX-LENGTH LINE FCK MLL / I, BRACKET FIGURE 28. LEADING EDGE DEVICE GAP, OVERHANG, AND DEFLECTION DEFINITIONS

deflections and grid positions are provided in Appendix F. The slats also

contained pressure instrumentation at four spanwise locations. The inboard

leading-edge slat could be replaced with a short-chord fixed-camber

Krueger. This FCK could be positioned at two deflection angles with two

grid positons at each angle.

The FCK did not contain pressure

instrumentation.

The trailing-edge high-lift system consisted of 80-percent span two-segment

flaps. The flaps were continuous, with no inboard aileron or exhaust gate.

They were installed in the desired positions using fixed brackets which

attached the main flap to the wing and the auxiliary flap to the main flap.

Each forward flap segment could he installed at four deflection angles, and

each aft flap segment could be installed at two deflection angles. The

bracket attachments were such that the aft flap angles were independent of

the forward flap angles, allowing either aft deflection and grid position to

be used with all four main flap settings. The exact flap deflections and

grid positions are given in Appendix F. The cruise configuration model

utilized the same flap linkage fairings as the cruise wing of the high-speed

LB-506A. For the flap-deflected case, a new set of fairings was used. The

new fairing were set in one position relative to the main flap, and

represented the fairing position for maximum fairing deflection. The

definitions of the flap gap and overhang are presented in Figure 29.

+ OVERHAN FIGURE 29. FLAP GAP, OVERHANG, AND DEFLECTION DEFINITIONS

The outboard ailerons on this model, attached with fixed brackets, could be

manually positioned at several deflection angles. The model was equipped

with inboard and outboard spoilers, as shown on the control surface diagram

of Figure 27. On the model, a one-piece bent-plate-type spoiler was used to

represent the airplane's three inboard panels, and a one piece

bent-plate-type spoiler was used to represent the outboard three panels.

A set of landing gear, which included two wing-mounted gear and one nose

gear, could be installed on the model for use in the landing or takeoff

configuration. The airplane gear wells and gear doors were simulated on the

model, and gear well fillers were provided for the gear-up case.

The horizontal and vertical stabilizers from the high-speed LB-506A model

were used on this model. The horizontal stabilizer was adapted to a

remote-drive and position-indication system, and was modified slightly to

match the new aft fuselage lines. The vertical fin was installed on this

model such that the exposed area was the same as on model LB-506A. This

placed the top of the vertical stabilizer at a different height due to the

change in aft fuselage lines. The dorsal fin was also used; however, the

contour of the dorsal was changed as shown in Figure 30. Horizontal and

vertical stabilizer diagrams are presented in Figures 31 and 32, respectively.

Two wing-mounted nacelles and pylons were used on this model. These parts

were the nacelle/pylon combination previously tested on model LB406A. The

flow-through nacelle represented that of the Pratt & Whitney Aircraft JTlOD

engine. The flap-linkage fairing incorporated into the pylon was modified

to allow the fairing to deflect with the flap.

LB-507A INSTRUMENTATION

The instrumentation associated with this model included a six-component

internal balance, wing static pressure orifices, a remotely driven

horizontal stabilizer, and an internal fuselage pitch system. The internal

pitch system and remotely driven horizontal stabilizer required the standard

Douglas power supplies, control console, and position readout systems. The

control console also included Douglas bubble-pack monitoring equipment.

MODEL LB-507A D 2A DORSAL L.E.

FUS (B,,) FIGURE 30. DORSAL FIN (DzA) MODEL LB-507A HORIZ STAB.

= 0.114 n? (1.2312 FT’)

sH

ORIGIN.

AR = 4.10 Y = 229.022 (90.166) h = 0.350

I-------

SWEEPC, = 30’ = lo.o” rH

:9”

- PIVOT AXIS 65.42% C Y = 245.209 (96.539s

1c

2 = 6.759 (2.661) THEORETICAL TRAILING EDGE L MODEL ACTUAL TRAILING EDGE (CUT BACK TO ACHIEVE 0.0254 cm DIMENSIONS IN CENTIMETERS (INCHES) MODEL SCALE (0.01 INCH) THICK TRAILING EDGE) FIGURE 31. HORIZONTAL STABILIZER (H ,D) MODEL LB-507A 12.057 - (4.7469) I -i = 0.086 rn2 (0.931 FT2) % = 1.600 AR h = 0.35 SWEEP C, = 35’ 217.998 (85.326) = 19.829 (7.807) 37.203 (14.647) = 15.617 (6.1483)

Y

+z” THEORETICAL t TRAILING EDGE ACTUAL MODEL :“U^T’-K~K~+ZE 0.0254 fo.01 I THICK) I VERT STAB. _/ ORIGIN DIMENSIONS IN CENTIMETERS (INCHES) MODEL SCALE UNLESS OTHERWISE SPECIFIED FIGURE 32. VERTICAL STABILIZER (V,D) Aerodynamic forces on the model were measured using the Ames Task Mark IIC,

10.16-cm (4-in.) diameter internal balance. The upper aft balance pin hole

was used for this installation.

Pressures over the model wing, aileron, and deflected high-lift system were

measured by 12 48-S-type scanivalves arranged in two 6-pat modules mounted

in the fuselage nose. Access to the scanivalves was obtained by removing

the nose and forward constant sections of the fuselage. In addition to the

four complete rows of pressure orifices, one pressure tap was located at the

trailing edge of an inboard station (18-percent semispan) to help evaluate

any separation that may have occurred (Figure 33).

The angle of attack of the fuselage reference plane was measured using a

bubble pack installed in the fuselage nose. From aFRP = -6O to Do,

the model was pitched using the external pitch system. From O" to +lO"

angle of attack, the fuselage was pitched using the fuselage internal pitch

drive while maintaining the balance at O".

For angles of attack +lO" to

+34O, the fuselage was pitched using the external pitch drive with a loo

angle maintained between the balance axis and the fuselage axis.

The horizontal stabilizer incorporated remote drive and a

A Douglas control panel and digital readout was

position-indication system.

provided for use in the tunnel control room.

LB-507A MODEL INSTALLATION

The model was mounted in the Ames 12-Foot Pressure Tunnel using the Douglas

tandem support system and the Ames Task Mark II 10.16-cm (4-in.) balance.

The balance was attached to the support struts using the Douglas balance

pitch block. The installation is depicted in Figure 34.

MODEL LB507A NOTES % b/2 1 WING PANEL

I I

18% i

I

I

(L) 72.5% M

I

(L)

I

I

I T.E. ONLY FIGURE 33. PRESSURE ROW LOCATIONS MODEL LB-507A OUTER HATCH 1 c INNER ACCESS HATCH \ / I I BAYONET-.

I

PITCH CONTROL / STRUT LENGTH -iDJUSTMENT

HATCH -H

-TURN+ABLE-/--- FIGURE 34. INSTALLATION IN NASA AMES 12-FOOT PRESSURE TUNNEL

LB-507A RESULTS AND DISCIISSIONS

Cruise Wing Characteristics

The initial configuration tested was the cruise wing body with the nacelles, pylons, and strakes attached.

The basic high-Reynolds-number

characteristics (lift, pitching moment, and drag) for the configuration are

shown in Figure 35.

Two different runs of the same configuration are shown

to indicate the repeatability'of the data. This figure indicates that a

tail-off CLMAX of 1.59 was obtained at the basic test condition of

M = 0.20 and RNMAC = 4.61 million.

This comparedwith a maximum value of

1.54 obtained from Phase I testing of the LB-486 model. A direct comparison

of the data from the two tests is shown in Figure 36. Besides a higher

the LB-507A model exhibited better tail-on pitching moments than

'LMAX'

did the LB-486 model. Though improved, the pitching moments of the LB-507A

model still included pitch-up prior to stall. Post-stall pitch-down was

abrupt and forceful.

-

Y MODEL LB-507A I- CONFIGURATION B 313W1BX1ElP1CN1CS11F =1 MACH = 0.20 0. 300 tl x IO= RN MAC = 4.61 e 5 0.200 e I: 5 0. 100 h I I .dem . -10 -5 RNGLE OF RTTRCK-DEE a -0. It00 El -0.500 m -0. boo -0. 700 q RNGLE OF RTTRCK-DEG J -0. 50 -0. 800 STC A. LIFT AND PITCHING MOMENT FIGURE 35. CRUISE WING CHARACTERISTICS AND REPEATABILITY Lt. 0 3. 5 : .

; J : , 3.0 L ; : : 2.5 I 2. 0 I. 5 I. 0 I- 0. 5 0. DRAG FIGURE 35. CRUISE WING CHARACTERISTICS AND REPEATABILITY MODEL LB-507 A AND MODEL LB-486 A MACH = 0.20 RN MAC = 4.61 x IO6 NACELLES, PY LONS, STRAKES ON

GEAR OFF 0

t 0.200

-I

v - RNGLE OF RTTRCK-DEG a I -0.200 V -0. 300 El q I 1 I I I IO I5 20 25 $ RNGLE OF ATTRCK-DEG B -0.502 -0.800 WC A. LIFT AND PITCHING MOMENT FIGURE 36. COMPARISON OF LB-507A AND LB-486B CRUISE WING CHARACTERISTICS MODEkLLE507A MODEL LB-466A Y. o- 3. 5- x L $ ‘3. o- L El 5 ‘2. 5- w V 0 QQO El B 2. o- Q v I. 5- I. o- SYM RUN TEST ‘H LB-507A OFF 0 3 0.5 I I I I 1 I I I I I I I I I I 0. 02 0. 0) 0. ob 0. 08 0. IO 0. I2 0. 19 0. lb 0.18 o.i.3 0. 22 0. a 0. 2b 0. 28 0. K, 0. 32 0.39 a DRIZ COEFFICIENT STC 0. DRAG FIGURE 36. COMPARISON OF LB-507A AND LB486A CRUISE WING CHARACTERISTICS

Mini-tuft pictures of the wing, for a Mach number of 0.20, are presented in

This figure

Figure 37 for angles of attack before and after C

LMAX'

illustrates the stall phenomena of this high-aspect-ratio wing at

= 4.61 milli.on. As was the case with the LB-486 model, the outboard

R%AC

wing panel stalled prior to the inboard panel. The inboard panel stalled

completely (separated to the leading edge) at an angle approximately 6O

Figure 38 presents the chordwise

higher than the outboard stall angle.

pressure distributions of the four streamwise pressure rows for aCLMAX

(13.61'), and lo and 3' past CYC suction peaks

At "'LMAX'

LMAX*

Slightly negative trailing-edge

are evident for all spanwise locations.

pressure coefficients are noted for this condition at all spanwise

Large spanwise flow angles are indicated in the corresponding

stations.

tuft photo for the trailing-edge region. AtaFRp = 14.59O (lo past

stall), the 72.5-percent semispan station plot indicates separation near the

= 16.54' (2O past stall), the 57-, 72.5-, and

leading edge. At aFRp

95-percent semispan stations are separated at the leading edge. On the

other hand, the inboard station was still heavily loaded.

The cruise wing configuration was

Reynolds number and Mach number effects.-

also tested at Mach = 0.20 at various reference chord Reynolds numbers,

ranging from 1.14 million (atmospheric conditions for the Ames facility) to

4.61 million. Test results are presented in Figure 39. Comparing the

results of the lowest Reynolds number run to the highest Reynolds number

was reduced from 1.59 to 1.31, o~C

data shows that C

LMAX was

LMAX

and the magnitude of the post-stall lift

reduced from 14.5' to 13.6',

loss is decreased. A positive CM shift was apparent for angles of attack

prior to stall, but the configuration still exhibited the same pitch

The maximum value of L/D was

variations for the angles just after C

LMAx'

reduced from 20.02 to 15.62 by the decrease in Reynolds number. Figure 39

suggests that C will not increase significantly, due to Reynolds

LMAX

number effects, as the Reynolds number is increased from the highest wind

tunnel value to flight conditions.

A. aFRP = 12.66” = 13.61° (a ” aFRP % ’ MAX FIGURE 37. MINI-TUFT PHOTOS-FOR CRUISE WING/EiODY WITH NACE’LLES (RUN 3) (CONTINUED) c. Q = 14.59O FRP D. aFRp = 16.64o FIGURE 37. MINI-TUFT ~0~0s FOR CRUISE wiNG/~00Y WITH NACELLES (CONTINUED) MODEL LB-507A PERCENT SEMISPAN = 35.09 PERCENT SEMISPAN = 57.00 -II- -II- .iYti RUN MACH ALPHA -12- -II- 0 =3 0.20 13.61 -II- -,I- -IO- -10.

-P- -9.

-I- -.- -,- -7- 0" -6- 0” -6.

-5.

-5- , -4- -4.

-3- -3- -2- -2- -I - -I- I : I I I L. II 20 !P 6 0 10 12 18 1) 22 d&D CHO; PERCENT SEMISPAN = 72.50 PERCENT SEMISPAN = 95.00 -II -I3- -12.

-II - -IO- -9- -8- -7.

u" -6.

-5.

-4- -3- -2- -1 - I ) I 24 18 26 CHORD A. aFRP = 13.61° (aqMAx) FIGURE 38. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING WITH NACELLES, PYLONS, AND STRAKES ATTACHED MODEL LBb07A PERCENT SEMISPAN = 35.09 PERCENT SEMISPAN = 57.00 -IS- -1X- SYM RUN MACH ALPHA -12- -12- 0 =3 0.20 14.59 -II - -Il- -ID- -IO- -6- -6- -I- -6- -,- -7- 0” -5- 0” -5- -5- -5- -a- -.- -1- -1- -2- -2- -I- -I - PERCENT SEMISPAN = 72.50 PERCENT SEMISPAN = 95.00 -137 -12- -II- -ID- -Q- -a- -7- 0” -6- -5- -4- -I- -2- -I - % I ( P = ‘14.59O B.

aFRP FIGURE 38. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING WITH NACELLES, PYLONS, AND STRAKES ATTACHED (CONTINUED) MODEL LB507A PERCENT SEMISPAN = 57.00 PERCENT SEMISPAN = 36.00 -13 -n- -12 SYM RUN MACH ALPHA -II- 0 =3 0.20 16.54 -II -II- -10 -,D- -0 -s- -I -B- -7 -I- 0” -5 & -5- -5 -5- -4 -t- -1 -Y.- -2 -2- -I -L- I h 2i

CHUG

PERCENT SEMISPAN = 95.00 PERCENT SEMISPAN = 72.50 -II- -1x- -12- -12- -II - -II- -10- -IO- -D- -S- -a- -B- -1- -7- & -s- 4 -5- -5- -5- -,- -4- -3- -3- -2- -2- -1 - -I- C.

aFRP = 16.54~ FIGURE 38. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING WITH NACELLES, PYLONS, AND STRAKES ATTACHED (CONCLUDED) MODEL LB-507A CONFIGURATION BgB W, B X,, P,= N,= MACH = 0.20 1.6 0.8 2 5 10 20 50 RN MAC ’ lo6 EFFECT OF REYNOLDS NUMBER ON MAXIMUM LIFT OF CRUISE WING FIGURE 39.

Figure 40 presents the influence of Mach number on the same configuration.

These data were obtained at a reference chord Reynolds number of 2.60

million.

The effect of Mach number was to decrease C

LMAX ICL~AX= 1044y

1.40, and 1.34 at Mach = 0.20, 0.26, and 0.32, respectively). Also,

increased Mach number tended to decrease the angle of attack for the

outboard stall.

Macelle/pylon/strake effects.- The effects of the nacelles, pylons, and

strakes are shown in Figure 41. Removal of the nacelles and pylons resulted

in a decrease in C from 1.59 to 1.47. The pitching-moment curves show

LMAX

the nacelles and pylons to be destabilizing prior to stall and stabilizing

after stall. The drag increment at l.2Vs due to the nacelles, pylons, and

strakes was 0.0171 and they reduced the L/D from 20.3 to 16.3. Mini-tuft

photos for the nacelles-off and pylons-off case are shown in Figure 42.

in local flow, compared to the configuration

Be1 ow CLMAXy improvements

with nacelles, were evident aft of the nacelle location. Outboard

separation patterns were similar for the nacelles on and off cases; however,

comparison of Figures 42 and 37 show that the presence of the nacelles

retarded flow separation on the wing region aft of the nacelles.

Chordwise pressure distributions for the configuration with the nacelles and

pylons removed are presented in Figure 43.

The angles of attack selected

are stall (11.55') and higher. At the aFR,, of 13.55', the

72.5-percent semispan station shows a collapse of the suction peak, while

the 95-percent semispan station shows only a modest increase in Cpmin and

mild trailing-edge separation. At a lo higher angle of attack, the

suction peak of the 57-percent semispan station collapsed. The most

outboard station remains reasonably well attached up to 16.5O angle of

attack, the same angle as the nacelles on case.

From the standpoint of low-speed clean-wing characteristics, the addition of

strakes to the nacelles is detrimental from both a lift and pitching-moment

standpoint. This detriment is illustrated in Figure 44. Addition of the

strakes reduced the tail-off clean-wing CLMAX from

1.62 to 1.59 and

increased the pre-stall nose-up moments.

we l&l Q 0 MODEL LB-507A I- CONFIGURATIONB W X P N S E 0 38 1B 1B 1C 1C 11F RN = 2.60 x 106 E MPC H 0.300- 0

2. 50

ii -0. 100-i RNGLE OF RTTRCK-DEG 2. oo- a s w B -0. 200 :: k I. 50- -0. 300 t !I I. oo- -0. YOO -0.500 -0. boo I I I I I I -5a * ) 5 IO I5 20 25

0 RNGLE OF RTTRCK-DEG

J

-0.50 -0. 800 STC A. LIFT AND PITCHING MOMENT W -4 FIGURE 40. EFFECT OF MACH NUMBER ON CRUISE WING w ru MODEL LB-507A q I. o- 0. 5 mh +oI . 4. I I I I I I I I I t I I I I 1 I 12 0. 00 0. 02 0. O$ 0. Ob 0. OR 0. IO 0. I2 0. IY 0. lb 0. I8 0. 20 0. 22 0. 2$ 0. 2b 0. 28 0. 30 0. 32 0. 3$ 0. 3b DRAG COEFFICIENT STC B. DRAG FIGURE 40. EFFECT OF MACH NUMBER ON CRUISE WING MODEL LB-507A CONFIGURATION B3B W,, XIB ki MACH = 0.20 z 0. 30( I- MAC = 4.61 x lo6 RN t u 3. 50 5 l- 0. 20c Y F 3.00 l- :: 0. IOC I3 l-l El 0 El (L Q -a u l-- I 1 “-88e I I I I I I 2. 50 -10 -5 * fi g u5 IO I5 20 25 30 u -o.ooc ,- I3 RNGLE OF RTTRCK-DEE 2. 00 % :: -0. 2oc I- # k $ I. 50 u I- I- -0. 300 I. 00 I- -0.YOO -0. 500 -0. boo I I 1 r I 5 IO I5 20 25 -0.700 RNGLE OF RTTRCK-DEG -0. 800 A. LIFT AND PITCHING MOMENT FIGURE 41. EFFECT OF NACELLES AND PYLON ON CRUISE WING MODEL LB-507A 0 0 2. 0 El I3 I3 El I. 5 T I I I I 1 1 I I I I I I I I I I - 0. 00 ,02 0. 02 0. OY 0. Ob 0. 08 0. IO 0. 12 0. IY 0. lb 0. IX 0. 20 0. 22 0. 2Y 0. 2b 0. 28 0. 30 0. 32 0. 39 DRRG COEFFICIENT STC B. DRAG FIGURE 41. EFFECT OF NACELLES AND PYLON ON CRUISE WING A. aFRp = 12.54’ FIGURE 42. MINI-TUFT PHOTOS FOR CRUISE WING/B~DY (RUN 113) MINI-TUFT PHOTOS FOR CRUISE WING/BODY (CONCLUDED) ‘%iGbRE-42.

MODEL LB-507A PERCENT SEMISPAN = 35.00 PERCENT SEMISPAN = 57.00 -1l- -II- SYM RUN MACH -12- -II- -II- -II- -10- -IO- -5- -s- -I- -.- -7- ?- 0” -5- E -c- -5- -5- -a- -4- -I- -,- -2- -1 - CHDRD PERCENT SEMISPAN = 95.00 PERCENT SEMISPAN = 72.60 -I3 -12 A.

apRP = 11.55’ (afgMAx) FIGURE 43. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING WITHOUT NACELLES AND PYLONS MODEL LB-507A PERCENT SEMISPAN = 35.00 PERCENT SEMISPAN = 57.00 -13- -13- SYM RUN MACH ALPHA -12- -12- 0 =113 0.20 13.55 -II- -II - -lO- -IO- -O- -P- -a- -5- -7- -7- 2 -6- u” -6- -5- -5- -,- -k- -3- -3- -2- -I - PERCENT SEMISPAN = 72.50 PERCENT SEMISPAN = 95.00 -13-l -I,- -12- -I,- -IO- -o- -I- -7- -1- -2- B. aFRP = 13.55O FIGURE 43. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING WITHOUT NACELLES AND PYLONS (CONTINUED) MODEL LB-507A PERCENT SEMISPAN = 57.00 PERCENT SEMISPAN = 35.00 -13- -Is7 t SYM RUN MACH ALPHA 1 -12- 0 =113 -II- 0.20 14.50 -II- -,,- -m- -IO- -6- -n- -6- -.- -7- -7- 0” -6- & -5- -5- -6- -‘- -1- -5- -I- -2- -2- -l- -I - ) ____ o- ____ o-------v---u P 1 1 I2 I4 I6 20 Cd PERCENT SEMISPAN = 95.00 PERCENT SEMISPAN = 72.60 -13- -13- -12- -II- -II - -II- -IO- -IO- -6- -6- -c- -c- -7- -7- 0” -6- 0” -b- -5, -,- -3- -2- D -.a-.- ,-+------ 1 p,/*- , I P 24 I6 I. 2d CHORD C.

aFRP = 14.50~ FIGURE 43. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING WITHOUT NACELLES AND PYLONS ATTACHED (CONTINUED) MODEL LB-507A PERCENT SEMISPAN = 36.00 PERCENT SEMISPAN = 57.00 -IS- -11- -12- -12- -II- -II - -*0- -lD- -9- -B- -I- -I- -7- -7- 0” -6- OQ -I- -5- -5- -a- -*- -1- -2- -L- 1 -__w __*’ - a-------- ---. .__- I , I.0 I P I. IS P I2 CHA PERCENT SEMISPAN = 72.50 PERCENT SEMISPAN = 95.00 -II ‘1 -12 -II -IO -D -I -7 & -6 -5 -4 -3 D. aFRP = 16.50° FIGURE 43. CHORDWISE PRESSURE DISTRIBUTIONS OF CRUISE WING WITHOUT NACELLES AND PYLONS (CONCLUDED) y88, w I n q CONFIGURATION B,, W,, X,B PIc N,C S,,, I- @ 0 El MACH = 0.20 5 El tl l-4 0.300- 0 k 0 jTJ0 RN = 4.61 x lo6 MAC El El I3 ki 3.50- El 2 0. 200-, 0 El 0: E I @El I: 3. oo- R 0. IOOJ .m E ati@ I?

B I =zeee , I ’ I I I 1 I ’ 2. !io- IO I5 20 25 30 -10 -5 :yl 1 5 / -0. 100 -I RNGLE OF RTTRCK-DEG 2. 00 m ; a # -0. 200 :: e i :: I. 50 u -0. YOO -0. 500, -0. bO0.

q Ge- I I I I I I -50 . 5 IO I5 20 25 -0. 700.

RNGLE OF RTTRCK-DEG -0. 50- -0. 800.

STC A. LIFT AND PITCHING MOMENT FIGURE 44. EFFECT OF STRAKES ON CRUISE MODEL LB-507A Lt. 0 3. 5 :: h K =I 3.0 k !

L 2 2.5 0 0 El El 2. 0 El El I. 5 I. 0 0. 5

I I I I I I

I I I I I I I I 1 I

0.09 0.Oh 0.ox 0. IO 0. 12 0. Ilt 0. lb 0. IX 0.20 0. 22 0. 2Y 0. 2b 0. 28 0. 30 0. 32 0. 3Y 0.'

DRRG COEFFICIENT STC B. DRAG FIGURE 44. EFFECT OF STRAKES ON CRUISE

Landing Configuration Characteristics

The primary landing configuration consisted of:

a two-segment flap deflected at Z"/lZo (main flap/auxiliary

1.

flap)

a slotted, leading edge, outboard slat deflected at 27'

2.

3. a slotted slat or short-chord FCK inboard.

The grid optimization studies for the inboard slat and inboard FCK are shown

in Figures 45 and 46, respectively. A comparison of the best slat position

versus the two best FCK positions is presented in Figure 47. The best

pitching-moment characteristics were those associated with the FCK deflected

at 70°. This configuration also resulted in the highest tail-off C

LMAX

Deflecting the FCK at 55' decreased the of the test, 3.08. C from

LMAX

3.08 to 2.94, decreased the stall angle from 17.2' to 15.Z", and

degraded the post-stall pitching moments. The inboard slat configuration

value between the two FCK values and exhibited the most

had a 'LMAX

undesirable pitching moment trends of the group.

Reynolds number and Mach effects.- The effect of Reynolds number on the

maximum-lift coefficient of the landing FCK configuration is shown in

Figure 48. Unlike the trends for the cruise wing, the trends for the

landing configuration suggest that the C of the landing configuration

LMAX

will increase beyond the wind tunnel values as the Reynolds number is

increased from the highest wind tunnel value to flight Reynolds number.

Any

effort to extrapolate the data to arrive at an estimated C

value for

LMAX

flight conditions would be unwise in light of the distinct break in the

versus Reynolds number curve for the cruise wing (Figure 39).

cL

MAX

The effect of Mach number on the maximum-lift coefficient for the same

landing configuration is depicted in Figure 49. Again the trends of the

cruise wing differed slightly from those of the landing configuration.

Whereas C of the cruise wing decreased monotonically with Mach number,

LMAX

of the landing configuration increased slightly as the Mach

the 'LMAX

number was increased from 0.20 to 0.26. As the Mach number was further

increased to 0.32, the C of the landing configuration decreased from

LMAX

2.88 to 2.79.

1 P

CONFIGURATION B ~AWl~XIBPICNICSllF MACH = 0.20 RN = 4.61 x IO6 MAC 6 = 25112.5 FLAP 3. 50.

3. co ,;A- A q v I I I I I IO 2. 50.

I52 2o 25 3o V El A El RNGLE OF RTTRCK-D&J I- 2. 00.

z =1 r L i I. 50.

3 .

t !!!

-0. 400 Q 1.00, Q INBOARD RUN 0. 50. SLAT -0. boo 12A e 9A $8 I El I 1 1 1 I 8B 5 IO 15 20 25 -5 * q go. A0 128

12c I

RNGLE OF RTTRCK-DEG -0. 50.

STC A. LIFT AND PITCHING MOMENT FIGURE 45. LANDING SLAT GRID OPTIMIZATION vV 9. 0 V MODEL LB-507A V A A V 8. 0 0 qfr OA v 7. 0

B

$ .

i

A

t: j b.0 : ;’ J : 5.0 * Y. 0 3. 0 OUTBOARD SLAT 2. 0 27A 27A 27A I. 0 27A 27A

-4

- I I I 1 I 8 I I I I I I I I I I i St 0 ‘. 0. 0 0. 04 0. 08 0. 12 0. lb 0. 20 0. 2Y 0. 28 0. 32 0. 3b 0. 90 0. LtY 0. 48 0. 52 0. 5b 0. b0 0. b’t 0. bX DRAG COEFFICIENT STC B. DRAG FIGURE 45. LANDING SLAT GRID OPTIMIZATION -.

MODEL LB-507A i CONFIGURATION B 3BW,BX1BN1CPIC811P # MACH = 0.20 0. 300, F = 4.61 x lo6 RN k MAC 6 = 25/12.5 B 3. 50 FLAP 6 = XXl27A 5 0. 200, LE !k g 3. 00 2. 50 0. 50 -0. bO0, I 1 I I , I 5 IO 20 -5 . 15 25 60. 700.

f! *

RNGLE OF RTTRCK-DEG -0. 50 -0. 800.

STC A. LIFT AND PITCHING MOMENT FIGURE 46. FCK OPTIMIZATION WITH LANDING FLAPS 9. 0 MODEL LB-507A 8. 0 -- I I I I I I I I , 8 I I I I I I IO 0. o+ 0. ox 0. I2 0. lb 0. 20 0. 2Lt 0. 2x 0. 32 0. 3b 0. YO 0. LtY 0. 98 0. 52 0. 5b 0. b0 0. b’t 0. b8 72 DRRG COEFFICIENT B. DRAG FIGURE 46. FCK OPTIMIZATION WITH LANDING FLAPS MODEL LB-607A CONFIGURATION B 3BW113XfBP1CNICS11F MACH = 0.20 RN MAC = 4.61 x lo6 = 25112.6 6 FLAP 3. 50 3. 00

w

2. 50

I?

I

I

m

ANGLE OF -0. 300- 0 -0. Ltoo- 2 @f@ -0.500- w” wo I I 5 IO RNGLE OF ~~TTRCK-DEG -0. 50 STC A. LIFT AND PITCHING MOMENT FIGURE 47. FCK AND SLAT COMPARISON WITH LANDING FLAPS 9. 0 MODEL LB-607A 8. 0 El q q 7. 0 El I i i kj b.0 5 5.0, v Y. 0, 3. 0, 2. 0 I. o- Q!l --ec,

I I I 1 I I I

I I I I I I 1 I I I

* 0.00 OY 0. OY 0. OS 0. 12 0. lb 0. 20 0. 2Y 0. 2X 0. 32 0. 3b 0. YO 0. LtY 0. Y8 0. 52 0. 54 0, b0 0. b’t 0. b8 0 72 DRRG COEFFICIENT STC B. DRAG FIGURE 47. FCK AND SLAT COMPARISON WITH LANDING FLAPS MODEL LB-507A CONFIGURATION B 313W1BX1BP1CN1C MACH = 0.20 6 = 70Al27A LE 6 = 25112.5 FLAP 3.2 2.8 2.6 2.4 2 10 20 50 RN x10 MAC FIGURE 48.

EFFECT OF REYNOLDS NUMBER ON MAXIMUM LIFT OF LANDING FCK/SLAT CONFIGURATION MODEL LB-507A CONFIGURATION B W 3B 1A ‘IS ‘IC NlC GIA RN = 2.89x106 MAC 6 LE = 70127A 6 = 25il2.5 FLAP TAIL OFF, GEAR DOWN I I I I I 2.4 0.20 0.22 0.24 0.26 0.28 0.30 0.32 MACH NUMBER FIGURE 49. EFFECT OF MACH NUMBER ON C LANDING FCK/SLAT CONFIGURATION LMAX

Figure 5J shows the effects of having the

Nacelles/pylons/strakes effect.-

nacelles, pylons, and strakes on the landing configuration with the slat

inboard. The nacelles and pylons had a degrading effect on the post-stall

pitching moments in that their addition eliminated the post-stall pitch-down

that was present (tail-off) with the nacelles and pylons off. The nacelles

and pylons had no significant effect on the maximum lift value for this

particular configuration.

The nacelle strakes, which were added to increase the C of the inboard

LMAX

slat configurations, were effective in that respect. The tail-on data of

Figure 51 showed that the strakes increased the tail-on CLMAX of the

inboard slat configuration from 2.94 to 3.08. As might be expected, the

strakes degraded the pitching-moment characteristics. Figure 52 shows that

the strakes had very little impact on the inboard FCK configuration.

MODEL LB-507A I- CONFIGURATION BsB W,, Xle B =1 MACH = 0.20 0.300- z RN = 4.61 x lo6 M’AC B 6 LE = lZCl27A s 3. 50 6 = 25i12.5 g FLAP 0. 200- it 0 0 l!!

3. 00 E0. IOO-

0 ?.I

0 0

El ti

Q

a 0 8

I

I 1 ve@J I I I 1

Q

I5 20 25 30 &IO q -5 * Q 5 IO 2. 50 El -0. IOO- El El RNGLE OF RTTRCK-DEE !i 2. 00 El q El E -0. zoo- q E 0 El I S I. 50

B

-0. 300

5 8

0 El

Jg B I. 00 -0. YOO- El El -0. 500- 0. 50 -0. bOO- I . ee @ I I I I I -5 0 5 IO 15 20 25 I -0. Eo- RNGLE OF ATTRCK-DEG R J -0. 50 -0.800’ STC A. LIFT AND PITCHING MOMENT FIGURE 50. EFFECT OF NACELLES AND PYLONS ON LANDING SLAT CONFIGURATION 9. c MODEL LB-507A

'1

8. C 7. 0 E El h El ; i $ b.0 L 3 5.0 .d

‘1

Y. 0 3. 0

2. 0 -I

I. 0

Yt0. 000. OLt 0. 08 0. 12 0. lb 0. 20 0. 2+ 0. 28 0. 32 0. 3b 0. YO 0. Yt

C. . 0. Y8 0. 52 0. 5b 0. b0 0. b’t 0. b8 (

DRRG COEFFICIENT STC B. DRAG FIGURE 50. EFFECT OF NACELLES AND PYLONS ON LANDING SLAT CONFIGURATION MODEL LB-507A s CONFIGURATION B W W 38 ,BX,BPtCN,CS,lFVIDHID =I MACH, = 0.20 t” 0.300- RN = 4.6’1 x lo6 MAC b 6 = lZCl27A e 3. 50 LE 3. 00 2 50 m Q RNGLE OF RTTRCK-DEG OQ m g -0. 200 I

A I

Q 1.00

B

0. 50 -0. bOO;l -5 9 5 IO 15 20 25 -0.7ooJ I RNGLE OF RTTRCK-DEG

J

-0. 50 -0. 800; STC J A. LIFT AND PITCHING MOMENT FIGURE 51. EFFECT OF STRAKES ON LANDING SLAT CONFIGURATION 9. 0 MODEL LB-507A 8. 0 El 7. 0 El b II .

i tl : b.0 i i : 3 5.0 I B Y. 0 3.0 2. 0, B I. 0, B B - I I I I I I I I I I I I I I I I I BY 0 IO 0. OY 0. 08 0. I2 0. lb 0. 20 0. 2+ 0. 28 0. 32 0. 3b 0. 40 0. w 0. 48 0. 52 0. 5b 0. b0 0. b’t 0. b8 I 72 DRRG COEFFICIENT STC B. DRAG FIGURE 51. EFFECT OF STRAKES ON LANDING SLAT CONFIGURATION -- +w MODEL LB-507A Li W CONFIGURATION B 38 lBXIBPICNICS1lFVIDHID k!

MACH = 0.20

; 0.300

RN = 4.61 x 106

MAC e

6 :: = 70Al27A 3.50 LE 0.200

ii

= 25/12.5 FLAP

F

z

3. 00, #q 0. 100 B

lil

i I I ~~ I I . I -10 I I -5 , c 5 IO 15 20 25 -0. 100-j

B

RNGLE OF RTTRCK-DEG

I

-0. 200 00. 300 q

L

-I B 0 1.00 I 0. 50 -0. bO0 I I I I I 1 -5 - -o.700 RNGLE OF RTTRCK-DEG -0. 50.

IO I5 20 5 -0. 800 STC A. LIFT AND PITCHING MOMENT FIGURE 52. EFFECT OF STRAKES ON LANDING FCK CONFIGURATION - v. 0 MODEL LB-507A B 8. 0 7. 0 I .

; rl : , b.0 i : : a : 5.0 r It. 0 3. 0 2. 0 I. 0 T- *.

N 0. 00 0. OY 0. 08 0. 12 0. lb 0. 20 0. 2Y 0. 28 0. 32 0. 3b 0. YO 0. YLt 0. YE 0. 52 0. 5h 0. ho 0. b’t o.bX 0 DRAG COEFFICIENT STC B. DRAG FIGURE 52. EFFECT OF STRAKES ON LANDING FCK CONFiGURATiON

Large inboard flap deflection effect.- In addition to testing the baseline

landing flap deflection of 25°/12.50, a deflection of 350/10° was

tested at two different grid positions. The original grid position included

a negative overhang of l%, and resulted in a slight reduction in C

LMAX

from that of the baseline deflection (Figure 53). Analysis of the mini-tuft

photos (Figure 54) and the trailing-edge press,ures (Figure 55) indicated

that the large deflection caused separation in the trailing-edge region. In

order to reduce the extent of trailing-edge separation, a new grid position

including a positive overhang of 1 percent was created by extending the

spoiler trailing edge. As the mini-tuft photos and trailing-edge pressures

show, the positive overhang was effective in reducing trailing-edge

separation problems. CLa = o increased by nearly 0.20 and C

LM.AX

increased compared to the baseline but only by 0.03.

The large deflection

did, however, result in a large drag increment at 1.3Vs (0.0405 and 0.0270

for the negative and positive overhang cases, respectively).

Takeoff Configuration Characteristics

Most of the work accomplished with takeoff configurations was directed

toward the use of sealed (zero gap) slats. The advantage of the sealed slat

The disadvantages are

is that it results in appreciably higher L/D values.

can result in poor stalling

that it provides lower values of CLMAX and

characteristics, particularly if a small amount of yaw is present at stall.

Figure 56 compares data for the slotted and sealed outboard slats,with an

FCK deflected at 55O inboard. The slat grid 20A was completely sealed,

the grid 208 had a small gap, and the grid 27A had a normal gap. As the gap

decreased from 2.55 to 2.40 and the

was decreased, the tail-off C

LMAX

pitching moments became more positive. The L/D values at 1.2Vs, on the

other hand, increased from 11.97 to 12.87. The mini-tuft photographs of

Figure 57 clearly show the earlier separation of the outboard panel for the

sealed slat configuration.

” MODEL LB-507A t CONFIGURATION B W 38 lBXIBPICNICSllF E MACH = 0.20 i 0.300 RN MAC = 4.61 x lo6 kl 6 LE = lZCl27A 3. 50 5 0.200- 0 L : 2 0. IOO- i? 0. ii n *-II0 15 . e6e r 5 I IO I I5 I - 2o I 25 1 1 30 0 a8 0 om 0 -0. IOO- RNELE OF ,TTR,%Ef 2. 00 ki

*q*

:: -0. zoo- El b I.90 I.5 -0. 300- go* -0. YOO- /y 0 8

0 w-

-0. 500

1 0 c

0. 50 r 38 I I I I I -5 * 5 IO I5 20 25 RNGLE OF RTTRCK-DEG

I El

-0. 50, -0. 800 I3 STC A. LIFT AND PITCHING MOMENT FIGURE 53. EFFECT OF LARGE FLAP DEFLECTION 9. 0, 8. 0, 7. 0, z .

r i r: b. 0.

L ; : 5. 0, ,s $. 0, 3. 0 2. 0 Q gQ El I. 0 - 4 I I I I I I I I 1 1 I I I I I I I OY 0.

00 0. o+ 0. ox 0. I2 0. lb 0. 20 0. 2Y 0. 28 0. 32 0. 3b 0. +o 0. YY 0. YX 0. 52 0. 5b 0. b0 0. bY 0. bX ( DRRE COEFFICIENT STC B. DRAG FIGURE 53. EFFECT OF LARGE FLAP DEFLECTION PAL8 E 19.12’, 2’ O.H. = 1% i”FRP A.

= 20.08”. 4- UCbllAX’ 0 FLAP DEFLECTION CONFlGURATlONS MINI-TUFT PHOTOS FOR 3s FIGURE 54.

MODEL LBBB7A CONFIGURATION B W fE IA ‘,A ‘1C NIC M = 020

-0.8 RN = 4.61 x lo6

MAC = 12Cl27A LE TAIL OFF t7 = 0.18 a.6 - CF TE

0 *

I I I I I ~-. mu 4 8 12 16 20 24 28 QFRP - DEG

qo

00 0

-0.8 - q = q = 0.35 0.35 -0.6 - cP TE

O Oo

-0.4 -0.4 -0 00 0 0 0 0 a 0 -0.2 0 -0.2 -0 000

0 q o

“FRP - DEG - DEG “FRP

L

a.2 a.2 - FIGURE 55. EFFECT OF LARGE FLAP DEFLECTION ON TRAILING EDGE PRESSURES FIGURE 55. EFFECT OF LARGE FLAP DEFLECTION ON TRAILING EDGE PRESSURES 1.12 MDDEL LB-507A -4.8 -0.8 q = 0.725 -0.6

CF

TE -0.4

0 q n

-0.2

qooo 0

3 0 0. 0 q

E!- I 0 I 3 ----kyko l2 l6 0 ok ffFRP - DEG FIGURE 55. EFFECT OF LARGE INBOARD FLAP DEFLECTION ON TRAILING EDGE PRESSURES (CONCLUDED) MODEL LB-507A 0 CONFIGURATION B W X lAPICNICSllF t 38 16 El 0 t” 0.300- 0 o 0 !!

MACH = 0.20 RNinAC = 4.61 x 106 ki 0 I3 6 = 5/10 Ei 3. 50 FLAP 000, 5 0.200- 800 A E 0 0 0 go f I3 ooo(Tj 0 m 6 0. IOO- A 3. 00 I E I!3 m ii!

.01 , 1. I 4 I I b 5 IO & 15 20 25 A 30 2. 50 go A a -0: IOO- ATTRCK-D:i? ‘13 ’ -0. 200- -0. bO0 I 1 I I I 5 IO I5 20 25 -0. 700 RNGLE OF RTTRCK-DEG -0. 50 STC A. LIFT AND PITCHING MOMENT FIGURE 56. OUTBOARD SLAT GRID OPTIMIZATION WITH FCK INBOARD I- 9. 0 MODEL LB-507A 8. 0 El Y. 0 m 3. 0 2. 0 27A I. 0 208 20A 20A lx?%* r?

---$;e I .- I I 1 I I I I I I I , I I I r- I 3Y c I. b 0 0. OY 0. 08 0. 12 0. lb 0. 20 0. 2Y 0. 28 0. 32 0. 3b 0. YO 0. wt 0. YX 0. 52 0. 5b 0. b0 0. bit 0. bX 0 DRRG COEFFICIENT STC B. DRAG FIGURE 55. OUTBOARD SLAT GRID OPTIMIZATION WITH FCK INBOARD A. SLOTTED SLAT OUTBOARD, aFRP = 20.94O B. SEALED SLAT OUTBOARD, aFRP = 20.B!Yi” FIGURE 57. MINI-TUFT PHOTO OF TAKEOFF CONFIGURATION SHOWING EFFECT OF OUTBOARD SEALED SLAT -116

Figure 58 compares the results of a sealed slat outboard with three

different inboard leading-edge configurations: a slotted FCK, a sealed

slat, and a clean leading edge. Because of the early stall of the inboard

wing not protected by a leading-edge device, the C

of the clean

LMAX

configuration was very low (2.09) and the pitching moments were very well

behaved. The C of the inboard sealed-slat configuration was 2.24

LMAX

while that of the slotted FCK was 2.40. The pitching-moment trends of the

FCK and the sealed slat were similar: both showed nose-down moments just

after stall, even in the absence of a tail.

The respective values of L/D at

1.2V, for the inboard clean leading edge, sealed slat, and FCK are 14.59, 13.50, and 13.57, respectively.

One concern with the sealed slats is that they can result in lateral

instability when stall occurs under a yawed condition. This tendency is

illustrated in Figure 59. With a sealed slat outboard, the inboard

sealed-slat configuration became laterally unstable at aFR,, = 19O; the

FCK at o~,-RR= 17.5'. However, with a slotted slat outboard, the

FCK/slat configurations remained laterally stable throughout the

angle-of-attack range investigated (Figure 60).

Strakes effects.- Figure 61 shows that the addition of nacelle strakes to

the takeoff configuration with sealed slats inboard and outboard caused only

The CLMAX increment

small changes in the lift and drag characteristics.

due to the strakes in conjunction with takeoff flaps and slats, 0.06, was

less than half that for the landing flaps and slats case, 0.14. As was the

case with clean wing and landing configurations, the strakes were

detrimental to the pitching-moment characteristics.

Figures.62 and 63 show the effect

Mach number and Reynolds number effects.-

of Mach number and Reynolds number, respectively, on the aerodynamic

characteristics of the takeoff configuration with an FCK inboard and a

As the Mach number was increased from 0.20 to 0.32, sealed slat outboard.

decreased from 2.20 to 2.15 and the pitching moments degraded

'LMAX

slightly. Below CLMAX, the drag polar was insensitive to Mach number.

versus Reynolds number curve of Figure 63 suggests that the

The %MAX

maximum lift coefficientwilI,continue to increase as the Reynolds number

increases towards the flight value.

_. .__ MODEL. LB-607A CONFIGURATION B 38 WIB ‘1B ‘1C NIC ‘11, :: E 0.300 MACH = 0.20 k RN = 4.61 x lo6 MAC w -0.500

I

cob0

1 1

INBOARD SYM RUN LE DEVICE 6 A. LIFT AND PITCHING MOMENT FIGURE 58. EFFECT OF INBOARD LEADING EDGE DEVICE WITH A SEALED SLAT OUTBOARD v. o-# MODEL LB-507A 8. o-

P

tt

n b. O- G I 5 5.0 -1 ” 0 0 Lt. 0 El El El El E El El El 3. 0 I3 El 2. 0, 1.0.

--e+ I I I 1 I I I , I I I I I I I I OY 0 10 0. OY 0. 08 0. 12 0. lb 0. 20 0. 2Y 0. 28 0. 32 0. 3b 0. YO 0. 99 0. $8 0. 52 0. 5b 0. b0 0. b’t 0. b8 C 72 DnR; COEFFICIENT STC B. DRAG FIGURE 58. EFFECT OF INBOARD LEADING EDGE DEVICE WITH A SEALED SLAT OUTBOARD MODEL LB-507A ;s CONFIGURATION B W 38 1B ‘1El ‘IC N,C HIA ‘1, MACH = 0.20 RN STABLE

t

(AC,)

A/3=-5’ I I I I -10 -5 0 30 5 10 15 (AC”) Afl=--5O -0.01 - I I I I I I I I -10 -5 0 5 10 15 20 25 30 ANGLE OF ATTACK (DEG) FIGURE 59. EFFECT OF SLATS ON ROLLING AND YAWING MOMENTS THROUGH STALL WITH SIDESLIP MODEL L0-507A CONFlGURATtON B 38 wlB ‘1, ‘IC NIC G,A MACH = 0.20 RN MAC = 4.61 x lo6 = 25/lie FLAP 0.08 STABLE t 0.06 - (AC,) Af3= -5O 0.04 - 0.02 - ‘La LMAX rC ---- -A-- I I I I I I -5 10 15 25 ANGLE OF ATTACK (DEG) -0.02 l- FIGURE 66.

EFFECT OF fNBOARD FCK SLAT DEFLECTION ON ROLLING MOMENT THROUGH STALL WITH SIDESLIP MODEL LB’607A e CONFIGURATION B W ,BxlBpICNICsllP”lDHtD 8 MACH = 0.20 z 0. 300- k = 4.61 x lo6 RNMAC 4: l-l 3. 50 6 = 6A/20A LE 6 = mo z 0. zoo- FLAP E . =o ‘H 3. 00 4 H 0. IOO- E RNELE OF RTTACK-DEG E- 2. oo- I5 tl tl Q kl I. 50-

t

!7

-0. 500- 0. 50 q El -0. -0. boo- 700- El q 0 0 0 El El STRAKES SYM RUN El ON -0.5oJ -0. 800’ OFF STC A. LIFT AND PITCHING MOMENT FIGURE 61. EFFECT OF STRAKES WITH TAKEOFF FLAPS AND SEALED SLATS 9. 0.

MODEL LB-507A 8. 0.

7. 0.

x - i : 1: b. 0.

L 3 5. 0.

I3 , El Y. 0.

3. 0.

2. 0.

DRRG COEFFICIENT STC B. DRAG FIGURE 61. EFFECT OF STRAKES WITH TAKEOFF FLAPS AND SEALED SLATS MODEL LB507A c- CONFIGURATION B W X N E 38 1B 1B lCplCsllF =I m w = 0. 300- RN 7 e 0 @ 0 0 MAC 2.60 x IO6 6 = 5/10 om FLAP :: 3.50- 6 = 55Al20A LE ‘z 0.200- G?

l!l 0 f, 0. IOO- 3. oo- I+ h h r 1 ,eee . I I I 1 1 I 0 5 IO I5 20 25 30 -10 -5 2.50- El -0. IOO- VELE OF RTTRCK-DEE -0. 200 @ -0. 300 Q b3 I .

e -0. uo0-l

l.OO- #y@

SYM RUN MACH ‘,,A,

-0. I300

I I I

-0. 700 -

015 -;; -0. zoo- STC A. LIFT AND PITCHING MOMENT FIGURE 62. EFFECT OF MACH NUMBER ON TAKEOFF FCK/SLAT CONFIGURATION 9. 0 MODEL LB-507A

go

3.o-

B

2. o- Q I. o- d d IQ 4.3 1 I I *, I I , I I I I I I I I I I ---- IY 0. 00 0. OY 0. OX 0. I2 0. lb 0. 20 0. 2Y 0. 28 0. 32 0. 33 0. YO 0. YLt 0. w 0. 52 0. 5b 0. b0 0. bY 0. I38 0 72 DRRG COEFFICIENT STC 6. DRAG FIGURE 62. EFFECT OF MACH NUMBER ON TAKEOFF FCK/SLAT CONFIGURATION MODEL LB507A CONFIGURATION B 3BW1BX1BP1CN1C MACH. = 0.20 6 = 55Ai20A LE = 6110 FLAP 2.8 2.6 2.4 2.0 i .a 1 2 5 10 20 RN MAC ’ lo6 FIGURE 63. EFFECT OF REYNOLDS NUMBER ON MAXIMUM LIFT OF TAKEOFF FCK/SLAT CONFIGURATION

Alternative Flap Settings.- In addition to the primary takeoff flap setting

of 5"/1OO, two other takeoff flap settings (O"/Oo and 150/10°)

were tested.

Figure 64 presents the basic aerodynamic characteristics for

the 15"/10° flap setting with a variety of leading-edge-device

combinations. The highest C was associated with the slat/slat

LMAX

configuration. The best pitching moment was associated with the FCK/slotted

slat configuration. The highest L/D values were associated with use of a

sealed slat outboard.

The basic aerodynamic characteristics of the aircraft with a clean trailing

edge are presented in Figure 65 for several leading-edge device

The combinations investigated included a sealed slat outboard

combinations.

with an FCK or sealed slat inboard, and a slotted slat outboard with a clean

This latter configuration was representative of an

leading edge inboard.

auto-slat system. Also shown are the characteristics of the cruise wing,

for reference. The pitching-moment curves show the obvious aerodynamic

benefit of an auto-slat system in improving stall behavior. Figure 66

summarizes the L/D values for the takeoff configurations.

Aileron and Spoiler Characteristics

Aileron effectiveness is presented for takeoff and landing configurations in

At pre-stall angles of attack, the aileron

Figures 67 and 68, respectively.

effectiveness was well behaved for most angles of attack, but near the stall

angle the effectiveness of the upward deflected aileron diminished. The

shape of the rolling moment curve with aileron deflection indicates, for all

flap settings, that the negative deflections (TEU) were more effective than

the positive deflections (TED). In many cases, the incremental rolling

moment obtained was more than twice as large as the corresponding value for

positive aileron deflection. (Good data for the landing flaps, with

positive aileron deflections are not available.)

MODEL LE507A + CONFIGURATION B W X 36 1B lBPICNICSllF MACH = 0.20 z 0. 300 tl = 4.61 x lo6 R%AC 6 = 15110

3. 50

FLAP 5 0. 100 3. 00 -I q og I I I I I 1 5 IO OkfJ 20 25 2. 50 I w CK-DEG 2. oo-

,Q

I. 50- Q q I o- B

e

0. 50- LE SYM RUN I LE DEVK :E I I I I I I I -5 * 0 5 IO I5 20 25 RNGLE OF RTTRCK-DEE -0. 50- STC A. LIFT AND PITCHING MOMENT FIGURE 54. AERODYNAMIC CHARACTERISTICS OF THE 15°/100 FLAP CONFIGURATIONS 9. 0 MODEL LB-507A 2. o- I. o- -6CI 0 1JI I I 1 I I I I I I I I I # I , 0. 3b 0. YO 0. LtY 0. YX 0. 52 0. 5b 0. b0 0. b4 0. b8 0 72 * 0. 00 OY 0. OLt 0. ox 0. 12 0. lb 0. 20 0. 2s 0. 28 0. 32 DRRE COEFFICIENT STC B. DRAG FIGURE 54. AERODYNAMIC CHARACTERISTICS OF THE 15°/100 FLAP CONFIGURATIONS MODEL LB-667A CONFIGURATION B W X 38 1B lBplC NICSllF MACH = 0.20 El RN = 4.61 x IO6 MAC 3.50- TAIL OFF q Itoo- 2. 50- RNGLE OF RTTRCK-DEG -0. 800- STC A. LIFT AND PITCHING MOMENT FIGURE 65. AERODYNAMIC CHARACTERISTICS OF CLEAN TRAILING EDGE CONFIGURATIONS - .

.-c _ MODEL LB-507A m j 2.57 D 0 I3 C I B B q O q q 2. 0’ I. 5- I. o- 0. 5- L @” hh 00 0 OQ A lTw3 I I I I I I I I I I I -ec; I I 1 TY 1 D2 0.00 0. 02 0. 09 0. Ob 0. 08 0. IO 0. I2 0. IY 0. lb 0. 18 0. 20 0. 22 0. 29 0. 2b 0. 2X 0. 30 0. 32 0.39 ( * DRRG COEFFICIENT STC B. DRAG FIGURE 65. AERODYNAMIC CHARACTERISTICS OF CLEAN TRAILING EDGE CONFIGURATIONS MODEL Lp-507A CDNFlGlJAAilON 6 MACH = 0.20 = 4.61 x lo6 RNhlAc 6 = o/o FLAP 6 = 5110 FLAP 6 6 FLAP LE DEVICE INBD,%TBD SYMBOL INBDlOUTBD MAIN/AUX 8°f200 SLAT*ISLAT= o/o. 5ilO. 15flO 5!i”1200 FCKlSLAT* o/o. 5/10.15/10 ---- j l * * 12Ol27.5’ SLAT/SLAT om. 5110. xv10 l -•-•-+ 55Oi27.5’ FCKISLAT 0l0.5/10.15/10 *SEALED SLAT 0.6 0.8 1 .o 1.2 1.4 1.6 1.8 2.0 0.4 cL FIGURE 66. TAKEOFF L/D SUMMARY MODEL LB-!iO7A CONFIGURATION B 3BW1BX1BPlCNlCVl~H~~ MACH = 0.20 RN = 4.61 x lo6 MAC = 8A/20A LE 6 = 5110 FLAP i, = 0 Q (DEG) TED b TEU -10 -20 aRH (DEG) FIGURE 67. ROLLING-MOMENT COEFFICIENT DUE TO AILERON DEFLECTION FOR SEALED SLAT TAKEOFF CONFIGURATION IilODEL LE507A CONFIGURATION B tBW~~XIBp~~N~~S~~~v~~H~~G~~ MACH = 0.20 = 4.61 x lo6 RNMAC 6 = 55Af27A LE = 25112.5 ‘FLAP iH = 0 0.020 - o( 0.01 6 - i E iL k 0.01: 2- s !z= z- s P o.oot % - i z 0.004 TED lb w TEU 20 10 -10 -20 aRH -0.008 FIGURE 68. ROLLING-MOMENT COEFFICIENT DUE TO AILERON DEFLECTION FOR THE FCK/SLAT ,; LANDING CONFIGURATION

Spoiler effectiveness for takeoff and landing configurations is presented in

Figures 69 and 70, respectively. The spoiler data indicated well-behaved

characteristics for both configurations, with increasing effectiveness shown

for increased flap deflections. The spoiler arrangement consisted of large

chord panels compatible with space available aft of the rear spar, and

spoiler span corresponding to flap span. This powerful spoiler

configuration was needed because of the reduced-roll-rate capability

associated with the high-aspect-ratio wings.

The effect of syrunetrical spoiler deflection with landing flap deflection is

shown in Figure 71. These results were obtained for out-of-ground-effect

conditions. The large spoiler chord and spanwise extent was very effective

in reducing the lift and increasing the drag; however, a significant

positive pitching-moment shift was also apparent. Mhile the reduction in

lift and increase in drag would result in greater deceleration on the

ground, the positive increment of pitching moment would tend to unload the

nose wheel. The ground effect on pitching moment, lift, and drag, with the

spoilers deflected, should be obtained in a future test program.

Landing Gear Effects

The effects of the landing gear are shown in Figure 72. The gear increased

CD by 0.0245 and decreased L/D at 1.3Vs (at CL = 1.864) from 11.55 to

9.92.

MODEL LBb07A CONFIGURATION B,, W,, XIB P,, N,, S,,, HID v,, MACH = 0.20 6 = 8AI20A LE = 5110 FLAP i, = O0 0.14 0.12 0.08 L

t

0.06

ii

P

-I i 0.04 a

,I

-0

0.02

ia

‘*

/’

I

I I I I

0 -5 -10 -15 -20 -25 -30 SPOILER DEFLECTION (DG) FIGURE 69. ROLLING-MOMENT COEFFICIENT DUE TO SPOILER DEFLECTION FOR THE SEALED SLAT TAKEOFF CONFIGURATION MODEL LB-507A CONFIGURATION Bgg W lBX,B pIc N,.S,,. “ID “ICI GIA MACH = 0.20 6 = 55A127A LE 6 = 25112.5 FLAP i, = O0 0.12 0.10

/

/

/

0.08

/

/

/

0.06 /

s

/

0.04

/

0.02 0 -5 -10 -15 -20 -25 -30 SPOILER DEFLECTION (DEG) FIGURE 70. ROLLING-MOMENT COEFFICIENT DUE TO SPOILER DEFLECTION FOR THE FCK/SLAT LANDING CONFIGURATION -. .-- MODEL LB-507A CONFIGURATION B 3BW1BX1BNlCP1CS11FGlA w G MACH = 0.20 H 0.300- RN = 4.61x lo6 MAC i 6 8 = 55Al27A

3. 50

LE 6 = 25112.5 g 0.200- FLAP Y

3. 00

2.50- 0 El 0 q El El -0. 100 RNGLE OF$TTRK-DEG + 2.00- 0 q B 0 El El a I3 -0.200 =I El tl 0 e El E 0 I. 50- t- El i-4 5 o q El I. oo- -0. YOO- Q o 0 a 0 El I3 El -0. 500- 0 0 0’0 0 0. !io- El El El -0. boo- El es’ I I I 1 I I -5 5 IO I5 20 25 i ’ [ -0.700 El RNGLE OF RTTRCK-DEG -0. 50- -0.800 STC A. LIFT AND PITCHING MOMENT FIGURE 71. EFFECT OF SYMMETRICAL SPOILER DEFLECTION I ) 0 9. ( MODEL LB-507A 0 I a Q j 7.c ie VI

Y. 0 a

a 3. 0 0 El 2. 0 El El SYM RUN 6SP El I3 q q El a cl= -ec!

I I I I I I > I 1 I I I I I I I I . ( * 0.00 Yt 0. OY 0. 08 0. 12 0. lb 0. 20 0. 2Lt 0. 28 0. 32 0. 37 0. YO 0. LtY 0. Y8 0. 52 0. 5b 0. b0 0. wt 0. bX 0, 72 DRRG COEFFICIENT STC i B. DRAG FIGURE 71. EFFECT OF SYMMETRICAL.SPOILER DEFLECTION MODEL LB-507A CONFIGURATION B 3BW1BX1BN1CP1CS11F Ii MACH = 0.20 =1 0. 300- RN = 4.61 x lo6 k MAC 3.50- 6 E = 70Ai27A LE 6 5 0. 200- = 25112.5 FLAP !2 E 0 -0. IOOl I RNGLE OF RTTRCK-Dz -0. 2007: -0. 300-l I. oo- -0. Ltoo- -0. 500- 0. 50- I . I I I I I i -5 0 5 IO 15 20 25 RNGLE OF RTTRCK-DEG -0. 50- -0. 800- STC A. LIFT AND PITCHING MOMENT FIGURE 72. EFFECT OF LANDING GEAR 9. 0, MODEL LB-507A El B El 0 I3 0 B 8. 0, 7. 0 : .

B 0 q ; : , Il.0 q : : 5.0 I v. 0 3. 0 2. 0 I3 I3 SYM RUN GEAR I- I. 0 , 8 :: :F;IF I3 a 1 I I 1 I I I I f I I I I --ed I I I f-l- I IO 0. OLt 0. 08 0. 12 0. lb 0. 20 0. 2* 0. 2x 0. 32 0. 3b 0. YO 0. VI 0. Ltx 0. 52 0. 5b 0. b0 0. t9t 0. bX 72 OLt c 1. c DRRG COEFFICIENT STC B. DRAG FIGURE 72. EFFECT OF LANDING GEAR

CONCLUSIONS AND RECOMMENDATIONS

Conclusions

As a result of wind tunnel testing conducted at the NASA Ames 12-Foot

Pressure Tunnel and the NASA Langley V/STOL Tunnel, the objectives set for

the EET Phase II investigation of high-lift systems for advanced transports

have been accomplished. This combined NASA/Douglas research effort has

demonstrated the aerodynamic benefits of advanced-technology high-lift

systems, has established a comprehensive data base for analysis of

developing methods, and has identified future development areas.

The following conclusions are drawn from the LB-486 data:

Reduced VCK deflections, compared to those employed during Phase I

1.

C

testing, provided no benefit in terms of additional

LMAXor

improved stalling characteristics.

2. With takeoff flaps, use of a sealed outboard slat with a clean

leading edge inboard provided significant improvement in L/D and

pitching-moment characteristics compared to the basic slat

This configuration resulted in a significant

configuration.

penalty in CLMAX. Use of an inboard sealed or small-gap slat at

an intermediate deflection is a candidate for future low-speed

testing.

The full-span FCK offered no obvious advantages.in high-lift

3.

performance compared to either a full span VCK or a full-span slat;

however, an FCK (especially a short-chord FCK) inboard, used in

conjunction with a slat outboard, provided the greatest improvement

in stalling behavior with only a relatively small loss in C

LMAX.

4. The revised slat-trim configurations tested showed less improvement

in pitching-moment characteristics and a larger loss in C

LMAX

than the short-chord FCK/slat (inboard/outboard) combination.

5. The use of a single-segment flaperon in place of the high-speed

without penalizing L/D or

aileron significantly increased C

LMAX

Replacement of the single-segment

pitching-moment characteristics.

flaperon with a two-segment flaperon resulted in an additional

small increment in maximum lift.

6. Comparison of aerodynamic data for equivalent configurations in the

Ames 12-Foot Pressure Tunnel and the Langley V/STOL Tunnel

indicated generally good agreement for the lift characteristics.

The comparisons indicated differences in pitching moment and drag.

The following conclusions are drawn from the LB-507 data: 1.

For the high Reynolds number test condition, the cruise wing

achieved a tail-off C of 1.59 and an L/D at 1.2V, of

LMAX

20.02. Pitch characteristics were influenced by changes in Mach

and Reynolds number.

2. The optimization of the leading-edge devices indicated superior

and pitching moments for the configurations with an inboard

CLMAX

FCK; the L/D values for the inboard sealed-slat and FCK

The sealed-slat configurations

configurations were equivalent.

exhibited lateral instability near stall under a yawed condition.

Improvement in aerodynamic performance and pitch characteristics

could result from further leading-edge-device optimization studies.

3. Testing of the highly deflected flap (35"/10") indicated little

increase in C but a large increment in drag.

LMAX'

4. Mach and Reynolds number effects were studied during the test

program for selected configurations. CL pitching moments,

MAX'

and L/D values tended to improve with increasing Reynolds number

and decreasing Mach number. Extrapolation of the wind tunnel data

to flight Reynolds numbers suggested further increases in maximum

lift are possible.

5. The nacelles and'pylons increased the cruise wing C

by 0.1;

LMAX

increment on the flaps-deflected configuration was

the 'LMAx

nearly zero. The presence of the nacelles and pylons tended to be

a post-stall stabilizing influence.

The strakes, which were added to improve the CLmax of the

6.

slatted configurations, were effective in that respect. The

additional CLmax for the inboard slat configuration with

landing flaps was 0.14; for the takeoff flaps, 0.06. The

strakes did not, on the other hand, increase the maximum lift

In

values of the cruise wing nor of the FCK configurations.

all cases, the strakes were detrimental to the longitudinal

stability.

Aileron effectiveness studies indicated that, for all flap

7.

settings,

negative deflections (trailing edge up) were more

effective than positive deflections (trailing edge down). In

some cases, the incremental rolling moment obtained with the

negative aileron deflections was more than twice that obtained

with the corresponding value for positive aileron deflection.

The effect of spoiler deflection on roll characteristics

8.

increased as flap deflection increases. Symmetrical spoiler

deflections for landing flap settings were very effective in

reducing lift and increasing drag.

Recommendations

Analysis of the Phase II study data has identified those areas where

continued work could result in further improvement of the technology. The

potential for improvement has been noted in the following low-speed

aerodynamic characteristics: pitching moments for high-lift configurations

and increases in maximum lift for both landing and takeoff configurations.

It is therefore recommended that future studies include the following:

1. The use of small gaps to improve the pitching-moment

characteristics of slat configurations without decreasing L/D.

2. The use of a slat that has a larger slot near the pylon than near

the fuselage, to increase the section CLMAX of the inboard wing

panel, and to promote a more rapid-inboard lift loss after stall.

3. Additional testing of the inboard short-chord FCK, in order to

increase the configuration L/D by reducing deflection and/or

closing the gap.

4. High-lift testing in ground effect at high Reynolds number.

5. Reduced landing slat deflections to increase C

LMAX'

6. Higher-Reynolds-number testing to determine CLMAX and

pitching-moment trends at conditions more closely matching those of

flight.

APPENDIX A LB-486 A,B,C CONFIGURATION NOTATION Simulates the DC-X-200 Model D-969N-21 fuselage. Full-scale B2A dimensions: Length = 42.29 m (138.8 ft); constant section diameter = 602 cm (237 in.). The aft fuselage tail cone uses The fuselage is configured for tandem the DC-10 model parts.

strut support system.

Simulates the DC-X-200 Model D-969N-21 wing and is lofted to '3B represent the airplane wing with a l-g load. Full scale dimensions: = 212.597 m2 (2288.457 ft2j; sW = 47.252 m (155.027 ft); aspect ratio = 10.502; bW = 0.1407; MAC = 5.351 m (17.555 ft). The model wing has a x removable leading edge, full-span VCK flap, trailing-edge two-segment flap, outboard aileron on one side, and spoilers.

The wing is constructed of Armco 17.4 steel and contains five rows of pressure orifices.

Wing-fuselage fillet for B2AW3B.

'2B Horizontal stabilizer for DC-X-200 (slab surface).

HIA Vertical stabilizer for DC-X-200 (slab surface).

"1A Flow-through, short core cowl nacelle configuration (2).

*2A New pylons for mating N2A to wing W2B (2).

'2A Nacelle strake configuration (attaches to N2A, 2 each '1A nacelle).

Main and nose landing gear defined for the DC-X-200'airplane.

GIA Main gear wheel wells with gear extended are not provided.

APPENDIX A (CONTINUED) The outboard aileron with inboard trim at Xl{ = 89.020 cm a2A (35.047 in.) and outboard trim at Xw = 109.480 cm (43.102 in.). The hingeline is located at 75% C.

Inboard spoiler segments fabricated as individual parts.

fl'f2 side, L Superscript R = right = left side, None = both sides.

f, and f2 inboard O" spoilers with sheet metal aft flA'f2A extension. Trailing-edge step is filled with wax and faired (LB-486A). This assembly was refurbished and the T.E. step filled with potting (LB-486C).

Outboard spoiler segments fabricated as one piece.

Leading-edge slat inboard of XGI = 36.367 cm (14.318 in.) and support at nominal gap = 2.25% C, D.H. = 2.0% C, and 6SLAT = 25'.

Leading-edge slat outboard of Xw = 36.367 cm (14.318 in.)

L2A and supported at nominal gap = 2.25% C, O.H. = 2.0% C, and ~SLAT = 35'.

Leading-edge variable-camber Krueger inboard of wing station L3A = 36.367 cm (14.318 in.) and supported at the nominal xw = 2.82% C, O.H. = -0.725% C, and 6vCK = 55O.

gap Leading-edge variable-camber Krueger outboard of wing station L4A = 36.367 cm (14.318 in.) and supported at the nominal xw = 3.5% C, O.H. = 1.0% C, and 6VCK = 55O.

gap The inboard VCK extension to the fuselage.

L5a APPENDI’X A (CONTINUED) The VCK section at the pylon interruption.

L6A Inboard main flap of a two-segment flap with inboard trim at FIA xw = 13.868 cm (5.460 in.) and outboard trim at = 30.793 cm (12.123 in.).

xW Inboard aft flap of a two-segment flap trimned to match FIA F2A and supported from FIA.

A single-slot flaperon with inboard trim at Xw = 30.793 cm F3A (12.123 in.) and outboard trim at XW = 43.411 cm (17.091 in.).

Outboard main flap of a two-segment flap with inboard trim at F4A = 43.411 cm (17.091 in.) and outboard trim at xw Xw = 89.020 cm (35.047 in.).

Outboard aft flap of a two-segment flap trimmed to match F4A F5A and supported from FqA.

Wing coordinates (spanwise, chordwise).

xw ’ yw Angle of attack, in degrees, of the fuselage reference plane CIFRP relative to the equivalent free airstream. Nose up is positive.

Aileron deflection, in degrees. Positive deflection is trailing edge down.

Aft flap deflection, in degrees (see Figure 51).

GFAFT Main flap deflection, in degrees (see Figure 51).

6F MAIN Slat deflection, in degrees (see Figure 481.

'SLAT

APPENDIX A (CONTINUED) VCK deflection, in degrees (see Figure 48).

'%CK Incidence angle, in degrees, of the horizontal stabilizer iH HIA Positive deflection is trailing edge down.

Sumnary Code Body + cruise wing.

sl B2AW3BX2Ba2A' BWXNPZLLFFFFF s2 2A 3B 2B 2A 2A JA 3A 4A JA 2A 3A 4A 5A

fJA, 2A, 3, 4, 5, 6’ Body+fJwed wQu+VCK

a2AfJ, 2, 2A leading-edge device+flaps+nacelles, pylons, and nacelle strakes +VCK filler blocks.

Configuration S2 - VCK filler blocks.

s3 S2-W3B+W3D' S2-W3B+W3D-fl,2 + flAfpA. Configuration s4 S3+inboard spoiler trailing-edge extensions.

B W X M P Z L L F F F F F 2A 3B 2B 2A 2A JA JA 2A JA 2A 3A 4A 5A s5 a2A flA, f2A, f3, f4, f5, f6. Body+flapped wing +slat and WUSS leading-edge+flaps+nacelles, pylons, and nacelle strakes.

APPENDIX B LB-486A,B,C DIMENSIONAL DATA COMPONENT UNITS MODEL SCALE FUSELAGE (B2A) .------ Length cm (in.) 198.77 (78.255) Maximum width cm (in.) 28.293 (11.139) Maximum height cm (in.)

28.293 (11.139) (w3B) WING Area m2 (ft2) 0.4696 (5.055) Span 2.221 (7.286) m (ft) Mean aerodynamic chord m (ft) 0.251 (0.825) Root chord (trapezoidal wing) cm (in.) 37.076 (14.597) Total root chord cm (in.) 51.895 (20.431) Tip chord (trapezoidal wing) cm (in.) 5.217 (2.054) Total tip chord cm (in.) 9.27 (3.65) Aspect ratio 10.502 Taper ratio 0.1407 Spanwise station of MAC cm (in.)

41.580 (16.370) Fuselage station of 25% MAC cm (in.) 160.28 (63.102) Sweepback of 25% Cw 28.57 deg Dihedral("lg") 4.5 deg HORIZONTAL STABILIZER (H, A) Area m2 (ft2) 0.1298 (1.397) Span cm (in.) 70.234 (27.651) MAC cm (in.) 19.91 (7.839) Root chord cm (in.) 27.384 (10.781)

APPENDIX 6 (CONTINUED) UNITS MODEL SCALE COMPONENT HORIZONTAL STABILIZER (H,A) (continued) Tip.chord cm (in.) 9.583 (3.773) . . Aspect ratio 3.800 Taper ratio 0.35 Sweepback of 25% chord : 30.0.

deg Dihedral 10.0 deg Fuselage station of 25% HMAC cm (in.) 247.36 (97.384) Tail length (25% WMACto 25% HMAC) cm (in.) 87.076 (34.282) VERTICAL STABILIZER ($A) --... -- Area m2 (ft2) 0.099 -.-(1 .060) Span cm (in.) 39.700 (15.630) cm (in.). 26.731 (10.524) MAC Root. chord cm (in.) 366759 (14.472) Tip chord cm (in.) 12.87 (5.065) -Aspect ratio 1.6 Taper ratio 0.35 Sweepback of 25% chord 35.0 deg Tail length(25% WMACto 25% VMAc) cm (in.) 82.301 (32.402) OUTBOARDAILEAR!! (azA) Area aft of hingeline ,, cm2 (in21 54.4 (8.44) Span % b/2 18.4 Chord aft of hingeline 25.0 % C,", SPOILER (fl,f2) cm2 (in") 47.2 Area (each) (7.32) cm (in.) 13.2 Span (each) (5.18)

APPENDIX B (CONCLUDED) COMPONENT UNITS MODEL SCALE SPOILER (f3,f4,f5,f6) Area (total, one side) cm2 (in') 104.660 (16.222) Span (total, one side) cm (in.)

43.835 (17.258) NACELLE (NzA) Length cm (in.)

32.00 (12.60) Maximum cowl height cm (in.)

13.7 (5.38) Inlet diameter (fan cowl) cm (in.)

9.85 (3.88) Exit area (gas generator) cm2 (in'.)

6.86 (1.06) Incidence of thrust line to FRP 1.6 deg Toe in 1.8 deg

APPENDIX C LB-486A,B,C GRID NOTATION SLAT GRID NOTATION All gaps and overhangs are percent of local wing chord Dimensions are model scale = 36.367 cm = 14.140 cm xW xw (14.138 in.)

(5.567 in.)

O.H.

GAP GAP O.H.

'SLAT -2.0 -2.0 2.25 2.25 25' 1.50 -1.0 -1.0 1.50 25O -2.0 -2.0 3.25 25O 3.25 -2.0 -2.0 2.25 15O 2.25 1.50 -1.0 -1.0 1.50 15O -2.0 -2.0 3.25 15O 3.25 +4.65 +7.54 = 0.0 5O = 0.0 -2.0 2.25 -2.0 35O 2.25 1.50 -1.0 1.50 -1.0 35O -2.0 -2.0 3.25 35O 3.25

APPENDIX C (continued) LB-486A,B,C GRID NOTATION SLAT GRID NOTATION All gaps and overhangs are percent of.local,wirig-chord Dimensions.are,model scale = 89.020'cm - =.36.367 cm xW

xw

(35.047 in.)

(14.138 in.)

O.H. NOTATION GAP O.H. GAP 'SLAT - -2.0 2.25 -2.0 25' 2.25 L2AD 1.50 -1.0 1.50 -1.0 25' L2AE -2.0 3.25 -2.0 25O 3.25 L2AF z +2.0 0 +2.0 zoo -N 0 L?AG

APPENDIX C (continued) LB-486A,B,C GRID NOTATION VCK GRID NOTATION All gaps and overhangs are percent of local wing chord Dimensions are model scale = 14.140 cm = 36.367 cm

xw $

(5.567 in.): (14.138 in.)

GAP O.H. GAP O.H.

'VCK 'VCK - - 55O 3.5 -1 51.31a" 2.82 -0.725 51.3180 2.82 -1.725 55O -2 3.5 -0.725 -1 51.318' 1.82 55O 2.5 51.3180 1.82 -0.275 55O 2.5 0 Xw = 36.367 cm = 111.274 cm

xw

(14.318 in.)

(43.809 in.)

-1 3.5 55" 3.5 -2 55" 55O -1 0 2.5 55O 2.5 = 14.140 cm Xl4 = 36.367 cm- xW (5.567 in.) (14.318 in.)

41.318O 2.82 -0.725 -1 45O 3.5 41.318' 0.82 -0.725 4o" 0.5 -1

APPENDIX C (continued) LB-486A,B,C GRID NOTATION VCK GRID NOTATION All gaps and overhangs are percent of local wing chord Dimensions are model scale = 36.367 cm = 14.140 cm xW xW (14.318 in.1 (5.567 in.)

NOTATION GAP O.H.

- - 'VCK 'VCK 1.82 -0.725 45O 2.5 -1 41.31a" L3AG 2.5 0 1.82 -0.725 45O 41.31 a0 L3AH xW = 111.274 = 36.367 cm xw (43.809 in.)

(14.318 in.)

3.5 -1 3.5 -1 45O 45O 2.5 -1 45O 2.5 -1 45O

APPENDIX C (Continued) LB-486A,B,C GRID NOTATIONS FCK GRID NOTATIONS All gaps and overhangs are percent of local wing chord Dimensions are model scale = 14.140 cm = 36.368 cm xW xW (5.567 in.) (14.318 in.)

GAP O.H. O.H.

GAP 'FCK 'FCK 31.065O 2.82 -0.725 35O 3.5 -1.0 31.065' 1.82 -0.725 35O 3.5 -1.0 31.065' 0.33 -0.33 35O 0.5 -0.5 35O 2.5 -1.0 35O 1.5 -1.0 35O 0.5 -0.5 35O 0.5 -0.5 31.065O 2.82 -0.725 45O 3.5 -1.0 31.065' 1.82 -0.725 45O 2.5 -1.0 31.065' 0.33 -0.33 45O 0.5 -0.5 = 36.368 cm = 111.036 cm xW xW (14.318 in.)

(43.715 in.)

45O 2.5 -1.0 45O 2.5 -1.0 L8AD 45O 1.5 -1.0 45O 1.5 -1.0 L8AE

APPENDIX C (CONTINIIED) LB-486A,B,C GRID NOTATIONS FCK GRID NOTATIONS All gaps and overhangs are percent of local wing chord Dimensions are model scale = 36.367 cm XW = 111.036 cm xW (14.318 in.) (43.715 in.)

GFCK GAP O.H. GFCK GAP O.H. NOTATION 45O 0.5 -0.5 45O 0.5 -0.5 51.065' 2.82 -0.725 55O 3.5 -1.0 51.065'. 55O 2.5 -1.0 1.82 -0.725) 51.065' 0.33 -0.33 55O 0.5 -0.5 = 36.368 cm XW = 111.036 cm xW (14.318 in.) (43.715 in.)

55O 2.5 -1.0 55O 2.5 -1.0 55O 1.5 -1.0 55O 1.5 -1.0 55O 0.5 -0.5 55O 0.5 -0.5 = 14.140 cm = 36.368 cm

xW xw

(5.567 in.) (14.318 in.)

5o" 0.05 -0.5 5o" 0.05 -0.5 L9AA 60' 0.05 -0.5 60' 0.05 -0.5 L9AB 7o" 0.05 -0.5 7o" 0.05 -0.5 L9AC

APPENDIX C (CONTINUED) LB-486A,B,C GRID NOTATION MAIN FLAP GRID NOTATION All gaps and overhangs are percent of local wing chord Dimensions tire model scale Inboard Flap and Flaperon Grid = 14 140 cm Xw = 43.411 cm

xw .*

(17.091 in.)

(5.567 in.)

GAP O.H. GAP O.H.

‘FMAIN 1.3 3.2 2.5 6.0 5O 0.8 3.2 1.5 6.0 0.8 2.2 1.5 4.0 1.3 2.2 2.5 4.0 1.6 1.1 3.0 2.0 15O 1.3 2.2 2.5 4.0 0.8 2.2 1.5 4.0 0.8 1.1 1.5 2.0 1.6 0.0 3.0 0.0 25' 1.3 0.0 2.5 0.d 1.3 0.5 2.5 1.0 0.8 0.5 1.5 1.0 1.9 1.1 3.5 -2.0 1.3 0.0 2.5 0.0 '350 1.3 0.5 2.5 1.0 1.i 0.5 2.0 1.0

APPENDIX C (Continued) LB-486A,B,C GRID NOTATIOH MAIN FLAP GRID NOTATIOM All gaps and overhangs are percent of local wing chord Dimensions are model scale OUTBOARDFLAP GRID = 89.020 cm = 43.411 cm xW

Xw

(35.047 in.)

(17.091 in.)

NOTATION O.H.

GAP ‘FMAIN 6.0 2.5 6.0 1.5 5O 4.0 1.5 4.0 2.5 2.0 3.0 4.0 2.5 15O 4.0 1.5 2.0 1.5 0.0 3.0 2.5 0.0 25' 1.0 2.5 1.0 1.5 -2.0 3.5 0.0 2.5 35O 1.0 2.5 1.0 2.0

APPENDIX C (Concluded) LB-486A,B,C GRID NOTATION AFT FLAP GRID NOTATION All gaps and overhangs are percent of local wing chord Dimensions are model scale FLAPERON DIFFERENTIAL POSITION = 43.411 cm xW (17.091 in.)

GAP O.H. NOTATION 25O 2.5 1.0 F3AR = 14.140 cm = 30.793 cm xW xw (5.567 in.) (12.123 in.)

GAP O.H. GAP O.H. NOTATION GFAFT 7.5O 0.3 0.8 0.4 1.1 loo 0.3 0.8 0.4 1.1 12.5' 0.4 0.4 0.5 0.5 15O 0.4 0.4 0.5 0.5 = 43.411 cm = 89.020 cm xW xw (17.091 in.) (35.047 in.)

GAP O.H. GAP O.H. HOTATION GFAFT 7.5O 0.5 1.5 0.5 1.5 loo 0.5 1.5 0.5 1.5 12.5' 0.75 0.75 0.75 0.75 15O 0.75 0.75 0.75 0.75

APPENDIX D LB-507A CONFIGURATION NOTATIONS Fuselage represents the ATMR-11 aft fuselage and center body.

'B3B - The fuselage nose is the same as the one used with fuselage The fuselage has cutouts for the tandem-strut-support B3A* system and wiper for horizontal tail. Fuselage length = 44.2492 m (145.9619 f-t). (F.S.), constant section diameter = 4.310 m (14.142 ft). (F.S.).

Flew technology wing, rigged to represent the airplane wing '1B - under a "lg" load at test conditions. Full scale trapezoidal dimensions: SW = 148.0 m2 ( 1600 ft'j; bW = 40.6198 m (133.267 ft.); AR = 11.10; x = 0.275; MAC = 4.054 m (13.300 ft); I? = 5O. The model has removable leading and trailing edges, spoilers, outboard ailerons, and four rows of pressure orifices.

Iding fuselage fillet for B3BH1B with two strut clearance '1B - holes added.

Inboakd conventional leading-edge slat extends from station L3A - X = 2.267 cm (5.758 in.) to Xw = 6.6464 cm (16.882 in.).

The slat extends in a streamwise direction and the inboard and outboard trims are streamwise. The inboard slat deflections are 8O and 12.5' (streamwise angle).

Outboard conventional leading-edge slat extends from L4A - = 6.943 cm (17.636 in.) to Xw = 17.532 cm xw (44.530 in.). The slat extends normal to the wing leading edge. The inboard trim is streamwise and the outboard is normal to the wing leading edge. The outboard slat deflections are 20' and 27.5' (streamwise angle).

APPENDIX D (CONTINUED) Inboard FCK with inboard trim normal to the wing leading edge L5A - 2.399 cm (6.093 in.).

at Xw = The outboard trim is such that the Krueger will seal against the pylon. The inboard FCK deflections are 55' and 70'.

Inboard main flap extends from station Xw = 1.8047 cm FIA - (4.584 in.) to xW = 6.883 cm (17.484 in.). The flap deflections are 5O, 15O, 25O, and 35'. A pressure row is located at Xw = 6.183 cm (15.704 in.) (left hand).

Inboard auxiliary flap trim station same as FIA. The F2A - deflection angles of the auxiliary flap are 10' and 12.5'.

The pressure row is located at Xw = 6.183 cm (15.704 in.).

Outboard main flap extends from station Xw = 6.895 cm F3A - (17.514 in.) to Xw = 14.059 cm (35.710 in.) at the flap leading edge and Xw = 14.133 cm (35.897 in.) at the flap trailing edge. The pressure rows are located at = 10.069 cm (25.575 in.) (left hand) and Xw = 12.807 cm xW (32.530 in.) (right hand).

The flap deflections are 5O, 15O, 25O, and 35'.

Outboard auxiliary flap trim station XW = 6.895 cm F4A - (17.514 in.) to Xw = 14.133 cm (35.895 in.).

The trim is streamwise aft to 30% C, at which point the cut alA - slants outboard to permit flap deflection. The aileron outboard trim station at the leading edge is

APPENDIX D (CONTINUED) = 17.661 cm (44.858 in.) and it is a streamwise cut. The xW aileron deflections available are -20°, -loo, O", +lO", and +2@. The aileron does not have a built in seal.

Wing flap linkage fairings representing D-3243-11 cruise bFIB - configuration from LB-506A. Four per side in.addition to the fairing incorporated into pylon.

bFlB deflected to maximum position to allow flaps to bFIC - deflect. One position only relative to the main flap.

One-piece bent plate representing the three inboard spoiler flA - segments having a 8.66 cm (22 in.) constant chord. (F.S.).

One-piece bent plate representing the three outboard segments f2A - having a 7.874 cm (20 in.) constant chord. (F.S.).

Main and nose landing gear defined for an EET/ACA airplane.

GIA - A 5.59% scale flow through nacelle representing the Pratt & NIC - Whitney JTlOD engine. This is the same nacelle configuration used with the LB-506A model.

A 5.59% scale pylon used in conjunction with the WIB wing plc - and the NIC nacelle. The pylon positions the nacelle centerline at +2O with respect to the FRP and toed-in 2O with respect to plane of symmetry. The pylon is the same one used in conjunction with WTM LB-506A.

APPENDrx D (CONCLUDED) Same dorsal profile as D,B with a modified leading edge D2A - contour.

LB-506 H,C horizontal stabilizer modified at inboard end to HID - match BgB fuselage. Remote control position capability.

S = 0.1144 m* (1.2312 ft2); AR = 4.10; x = 0.350; sweep CV,4 = 30°; r = lO.OO.

LB-506 V,c vertical stabilizer modified at the root to match '1D - V3B fuselage. Sv = 0.0865 m2 (0.9312 ft2); = 1.600; x = 0.35; sweep Cv,4 = 35O.

AR Nacelle strakes from DC-10 model LB-246 on Nlc nacelle.

'1lF -

APPENDIX E LB-507A DIMENSIONAL DATA COMPONENT UNITS MODEL SCALE FUSELAGE B3B Length cm (in.) 248.680 (97.908) - Constant Section (9.480) Diameter cm (in.) 24.079 W,b WING Trapezoidal gross area m2 (ft2) .4645 (4.9997) Sweepback of the quarter chord 26.00 deg Taper ratio 0.275 Aspect ratio 11.10 Trapezoidal root chord cm (in.) 32.090 (12.636) 8.840 (3.480) Tip chord cm (in.)

Mean aerodynamic 0.2266 (0.743) m (ft) 2.271 (7.449) Span m (ft) Spanwise location of MACW cm (in.) 46.007 (18.113) Dihedral (lg) 5.00 deg VERTICAL STABILIZER V,D Gross area m2 (ft2) 0.086 (0.931) Aspect ratio 1 .6 Taper ratio 0.35 Sweepback at c/4 35.0 deg Theoretical root chord 34.442 (13.560) cm (in.)

Theoretical tip chord cm (in.) 12.070 (4.752) Mean aerodynamic chord cm (in.) 25.054 (9.864) Spanwise MACV position cm (in.) 15.616 (6.148) Horizontal distance cm (in.) 100.952 (39.745) from 25% cW to 25% c -V NOTE: All dimensions listed are in the FRP system.

All angles listed are in the WRP system.

APPENDIX E (continued) UNITS MODEL SCALE COMPONENT HORIZONTAL STABILIZER HID (1.231) Gross area m2 (ft2) 0.1144 Aspect ratio 4.10 0.35 Taper ratio 30.0 Sweepback at c/4 dw cm (in.) 68.4886 (26.964) Span 24.750 (9.744) Theoretical root chord cm (in.)

cm (in.) 8.656 (3.408) Theoretical tip chord Mean aerodynamic chord cm (in.) 17.983 (7.080) cm (in.) 14.371 (5.658) Spanwise MACH position Fuselage station of (0.25)MACH cm (in.) 243.507 (95.869) 10.0 Dihedral angle deg cm (in.) 124.419 (48.984) Horizontal distance from 25% cW to 25% cH OUTBOARDAILERON (a,A) 25.0 Chord aft of hinge line %$, cm (in.) 22.793 (8.974) Span INBOARD SPOILER (f,A) Area cm2 (in2) 3.027 (1.192) Span cm (in.) 29.538 (11.629) (1.230) Chord cm (in.) 3.124 OUTBOARDSPOILER (f2A) cm2 (in2) 3.453 (1.360) Area Span cm (in.) 37.051 (14.587) cm (in.) 2.841 (1.118) Chord NOTE: All dimensions listed are in the FRP system.

All angles listed are in the WRP system.

APPENDIX E (CON'TINUED) COMPONENT MODEL SCALE UNITS NACELLE (N,$ length cm (in.) 30.526 (12.018) Maximum cowl height cm (in.) 15.728 (6.192) Inlet diameter (fan cowl) cm (in.) 9.327 (3.672) Inlet area (fan cowl) cm2 (in') 68.284 (10.584) Exit area (gas generator) cm2 (in2) 16.258 (2.520) Incidence of thrust line to FRP 2.0 de3 Toe in 2.0 deg LEADING-EDGE SLAT (L3A, L4A) _----_--__ - Inboard) cm (in.) 27.150 (10.690) span (LsA - Outboard) cm (in.) 68.199 (26.850) Span (L4* Effective span %b/2 82.476 INBOARD-MAIN FLAP (FIA) Area cm2 (in2) 170.291 (26.395) Span cm (in.) 33.329 (13.122) Root chord cm (in.) 5.386 (2.120) Tip chord cm (in.) 4.837 (1.904) Inboard trim (X,) cm (in.) 11.643 (4.584) Outboard trim (XW) cm (in.) 44.409 (17.484) INBOARD-AUXILIARY FLAP (FZA) Area cm2 (in2) 105.631 (16.373) Span cm (in.) 32.766 (12.900) Root chord cm (in.) 3.225 (1.270) Tip chord cm (in.) 3.225 (1.270) Inboard trim (X,) cm (in.) 11.643 (4.584) Outboard trim (X,) cm (in.) 44.409 (17.484) NOTE: All dimensions listed are in the FRP system.

All angles listed are in the WRP system.

APPENDIX E (CONCLUDED) MODEL SCALE UNITS COMPONENT OUTBOARD-MAIN FLAP (F3A) 181.997 (28.210) cm2 (in21 Area 46.217 (18.196) cm (in.)

Span (1.902) 4.831 Root chord cm (in.)

(1.187) cm (in.) 3.014 Tip chord 210.168 (17.514) cm (in.)

Inboard trim (X.1 90.980 (35.819) cm (in.)

Outboard trim (X,1 OUTBOARD-AUXILIARY FLAP (F4A) 121.052 (18.763) cm2 (in21 Area 46.689 (18.382) cm (in.)

Span (1.230) 3.124 Root chord cm (in.)

(0.804) cm (in.) 2.042 Tip chord 44.485 (17.514) cm (in.)

Inboard trim (X,1 cm (in.) 91.403 (35.985) Outboard trim (X,1 NOTE: All dimensions listed are in the FRP system.

All angles listed are in the WRP system.

APPENDIX F GRID NOTATION LB-507A SLAT GRID NOTATION All gaps and overhang are percent of local wing chord Dimensions are model scale = 11.117 cm = 44.447 cm xw xw (4.377 in.)

(17.499 in.)

GAP O.H. GAP O.H.

'SLAT Inboard 6.00 8A 0.15 6.00 0.30 8B 0.65 6.00 0.80 6.00 0.46 4.00 12A 0.53 4.00 12.5A 1.50 -1.00 1.50 -1.00 = 44.447 cm = 91.173 cm

xw xw

(17.499 in.) (35.895 in.)

Outboard 0.00 2.00 0.00 2.00 20A 20B 0.50 2.00 0.50 2.00 27.5 2.25 -2.00 2.25 -2.00 -1.00 27.5B 1.5 -1 .oo 1.50

APPENDIX F (CONTINUED) GRID NOTATION LB-507A FCK INBOARD GRID NOTATION All gaps and overhang are percent of local wing chord Dimensions are model scale = 11.117 cm = 44.447 cm xw xW (4.377 in.)

(17.499 in.)

GAP OVERHANG 'SLAT 55A 0.75 -0.75 55B 1.50 -1.0 70A 0.75 -0.75 70B 1.5 -1 .o

APPENDIX F (CONCLUDED) GRID NOTATION LB-507A INBOARD TWO-SEGMENT FLAP (F,A/F2A) All gaps and overhangs are percent of local wing chord Dimensions are model scale = 44.447cm xW (17.449 in.)

MAIN AUX OVERHANG GAP GAP OVERHANG 'FMAIN 6FAUX - 5.0 1.50 4.0 10.0 0.50 1.50 15.0 1.50 2.00 10.0 0.50 1.50 25 2.50 0.00 12.5 0.75 0.75 35.0 2.50 -1.00 12.5 0.75 0.75 The inboard flap is rigged at the above station, and at the side of fuselage = 11.117 cm, (4.337 in.) with the same physical gap and overhang.

xw The outboard flap is also rigged to the above percent gap and overhang values at station Xw = 44.447 cm (17.499 in.). At all stations, outboard, the gap and overhang are the same percentages of the local wing chord.

REFERENCES 1. Oliver, Wayne R.: Results of Design Studies and Wind Tunnel Tests of an Advanced High Lift System for an Energy Efficient Transport. NASA CR-159389, 1980.

2. Steckel, Doris K.; Dahlin, John A.; and Henne, Preston A.: Results of Design Studies and Wind Tunnel Tests of High Aspect Ratio Supercritical Wings for an Energy Efficient Transport. NASA CR-159332, October 1980.

3. Crowder, J.P.: Fluorescent Mini-Tufts for Non-Intrusive Flow Visualization. McDonnell Douglas Report MDGJ7374, February 1977.

4. The Staff of Douglas Aircraft Company: Selected Advanced Aerodynmamic and Active Control Concepts Development. Summary Report. NASA CR-3469, 1981.

2. Govanmmt Accdm No.

1. Rqmrt No. 3.

NASA CR-3523 4. Tii md Subtitle 6. Raport Date July 1982

WIND TUNNEL TESTS OF HIGH-LIFT SYSTEMS FOR ADVANCED

TRANSPORTS USINGHIGH-ASPECT-RATIO SUPERCRITICAL WINGS

'* RrfwminOOr~nir8timwr

8. &forming Orgsniution Report No.

7.kthorw John B. Allen, Wayne R. Oliver, and

ACEE-17-FR-1608

Lee A. Spacht 10. Wak Unit No.

8. krfaming Organization Name and Address Douglas Aircraft Company 11. Contract or Grant No.

McDonnell Douglas Corporation 3855 Lakewood Boulevard NASl-15327 Long Beach, CA 90846 13. Type of Report nd fkiod Covered 12. Sponsoring Agency Name and Address Contractor Report National Aeronautics and Space Administration 14. Sponsoring Agency Coda Washington, D.C. 20546 15. Supphnentary Notes Langley Technical Monitor: Thomas G. Gainer Final Report 16. Abmact The wind tunnel testing of an advanced-technology high-lift system for a wide body and a narrow body transport incorporating high-aspect-ratio supercritical wings is described. This testing has added to the very limited low-speed high-Reynolds-number data base for this class of aircraft.

The experimental results included the effects on low-speed aerodynamic characteristics of various leading- and trailing-edge devices, nacelles and pylons, ailerons, and spoilers, and the effects of Mach and Reynolds numbers.

7. Key Words (Suggested by Author(s)) 18. Distribution Statement dings iigh aspect ratio High lift devices Supercritical Lift FEDD Distribution -ift augmentation Pitching moments Subject Category 02 iigh Lift Flow visualization 22. Rice 19. Security Qassif. lof this report) 21. No. of P-s 20. Security Classif. (of this page) Unclassified Unclassified 188 Available: NASA’s Industrial Applications Centers NASA-Langley, 1982