NASA TECHNICAL MEMORANDUM

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WIND TUMYEL INVESIlGA7'ION OF A LARGESCALE UPPER SL'RFACE

BUIWN- MODEL HAVING FOUR ENGINES

Kiyoshi Aoyagi, Michrel D. Palm& and David G. Koenig

Amcs Re euch Center

U. S. Amy Air Mobility IUD hbontory Moffett Field, Crlif. 94035

July 1975

BE boundary layer control c chord measured parallel to the plane of symetry, a (ft) pressure coefficient. Pt PSI% C~ - c t horizontal tail chord measured parallel to the plane of syletry, r (ft) E mean aerodynamic chord of wing, 2fS r (ft) CD drag coefficient, drag/Q ram drag coeificient, W/gc C~ram c~ jet moment lo coefficient , Fg/gS C~ lift coefficient, lift/~S Cz rolling-moment coefficient about stability axis, rolling .orent/%Sb Cm pitching-moment coefficient about 0.40 c, pitching nnent/&~S yawing-noaent coefficient about stability axis, yawing .orent/&~Sb C waentum coeificient W/gqaDS M C side-force coefficient about stability axis, side force/qaoS Y F~ static (wind off) incremental axial force due to flap deflection with power on, N (lb) F gross thrust with engine alone, N (Ib) (obtained statically) 8 static (wind off) incremental normal force due to flap deflection with F~ power on, N (lb) resultant force N (lb) Fh Tz, g acceleration of gravity, 9.81 m/sec2 (32.2 ft/sec2) i horizontal tail incidence, deg t LE z local static presure, I+i/w2 (lb/sq ft)

iii

pRFX:mING PAGE Ps free-stream static pressure. ~/m~(lb/sq ft) free-stream dynamic pressure, ~Irn~(lblsq ft)

S wing area, m2 (sq ft)

V free-stream air velocity, m/sec (ft/sec) or velocity based on isentropic expansion

W engine inlet weight rate of flow, kglsec (lb/sec) or weight rate of flow at blowing nozzle

WCP wing chord plane x chordwise distance from wing leading edge, cm

Y spanwise distance perpendicular to the plane of symmetry, m (ft]

lower surface distance frm UCP, cm YL Yu xpper surface distance from WCP, cm a of , deg

8 sideslip, deg

deflection, deg 6ai1

6f deflection of Coanda plate measured parallel to the plane of symmetry, deg (see fig. 2(f))

6 trailing-edge second flap deflection measured Farallel to the plane of f2 syetr:, deg (see fig. 2(f))

6 jet exhaust deflection angle wing off, tan-' FN/FA, deg (average value) j 6 slat deflection, measured parallel to the plane of symetry, deg S rl spanwise extent, y/ (b/ 2)

flap system stati' turning efficiency, F /F (average value) "f Rg wing leading edge sweep, deg A~~ Subscripts : ail aileron

KAC LE leading edge u uncorrected WIND TUNNEL INVESTIGATION OC A LARGE-SCALE

UPPER SURFACE BLDUN-FLAP MODEL HAVING FOUR ENGINES

Kiyoshi Aoyagi*, Michael D. Falarski**, and David G. Koenig*

*hes Research Center and **U. S. A-f Air Mobility R&D Laboratory

Investigations were conducted in the Ames 40- by 80-Foot Wind Tunnel to deteraine the aerodynamic characteristics of a large-scale subsonic jet transport model with an upper surface system. The model had a 2S0 of tspect ratio 7.28 and four turbofan engines. The lift of the flap system was augmented by turning the turbofan exhaust over the Coanda surface. Results were obtained for several flap deflections with several wing leading-edge configurations at jet mumentun coefficients from 0 to 4.0.

Three-colponcnt longitudinal data are presented with four engines operating. In addition, longitudinal and lateral data are presented with an engine out.

The maximum lif'. and angle of the four engine model were lower than those obtained with a two engine model that was previously investigated. The addition of the outboard had an adverse effect on tnese values. Efforts to improve these values were successful. A maximum lift of 8.8 at an angle-of-attack of 27O was obtained with a jet thrust coefficient of 2 for the landing flap configuration.

Lift augmentation by the upper surface blown-flap (USB) concept is currently being considered in some powered- 1ift transport designs. An im- portant factor for this consideration is the noise reduction due to wing shielding to a ground observer during the takeoff and landing operation of an upper surface blowing .

A wind tunnel investigation of this concept with a large-scale 2S0 swept -wing transport model having two engines has been reported in references 1 and 2 for aerodynamic and noise characteristics, respectively.

In order to determine the effects of four engines and increased spanwise extent of the Coanda surface on the aerodynamic characteristics of a large- scale USB transport model, two nacelles were added outboard of the existing nacelles on the wdel reported in reference 1. The resulting four engine configuration was investigated in the bes 40- by 80-Foot Wind 'humel. Aero- dynamic and noise characteristics of the model were obtained with several flap deflections and leading-edge configurations at jet roerent- coefficients from 0 to 4.0. O?11**the lerodynaric characteristics of the model will be presented in this report.

This report presents basic data of two wind tunnel investigations. The first investigation determined the aerodynamic characteristics of the model with the wing leading edge cmpletely swept and then unswept from the outboard nacelle to thz fuselage. This modification was made to improve tho maxirue lift and the stall angle of the rodel. The second investigation was made to determine the aerodynamic characteristics of the model with improved leading-edge devices and with BLC along the unswept leading-edge section and along the sides of the nacelles. The data with the horizontal tail on were obtained only during this investigation. The data of both investigations were obtained at Reynolds numbers from 2.1 x lo6 to 3.0 x lo6, based on a mean aerodynamic cbrd of 1.69 m (5.56 ft) and at dynamic pressures from 239 to 479 It/m2 (5 to lopsf), respectively.

MODEL AND APPARATUS

Twwind tunnel investigations were undertaken with the model. For the first investigation (Test 434), the wing leading edge was swept as shown in figure 1 (: Lazer, during the same investigation, the wing leading edge was unswept to O0 between the nacelles and between the inboard nacelle and the fuselage as shcwn in figure l(b). For the second investigation (Test 441), BE nozzles were added to the unswept leading edge and along the sides of the nacelles. In addition, highly cambered slats were installed at these sections.

Pertinent dimensions of the model are given in flgure 2(a). This model has the same geoinetry as that reported in reference 1 except as follows: the wing airfoil sections were altered from a NACA 63 series to a modified super- critical section, and the wing thicknesses were increased from 0.14~ to O.1Sc and 0.11~ to 0.12~at the root and tip respectively; the aileron was extended inboard from n = 0.75 to 0.70; the outboard nacelles were installed at rl = 0.48; and the Coanda surface was extended out to the aileron. Wing

The wing had a quarter chord sweep of 2S0, an aspect ratio of 7.28, and an incidence of 0'. The airfoil had a modified supercritical section that was .15c thick at the root and .12c thick at the tip. The ordinates of these sections are given in Table I. The wing tapered linearly in thickness between these two sections. For the first wind tunnel investigation, the entire wing leading edge Was swept. Later, the wing leading edge was unswept to O0 for the wing extendirg from n = 0.087 to 0.190 and frw s = 0.321 to 0.413 because of the leading edge flow separation problem at these sections. This was ac- complished by adding a chord extension to the existing swept wing leading edge as shown in figure 2(b).

Leading-edge devices

Figure Z(c) shows the l~adingedge configuratioxs used during the first wind tunnel investigation. When the wing leading edge was fully swept during the first wind tunnel investigation, a 0.15~ slat was deflected 60° with a 0.01% gap from n = 0.087 to 0.190 and n = 0.326 to 0.413, and a 0.25~ slat was deflectcd 52O from q = 0.546 to 1.00. The 0.15~ slat was also used as a deflected 68'. Wen the wing leading edge was lmswept from n = 0.087 to 0.190 and s = 0.326 to 0.413, a constant 0.2410 r (-79 ft) slat was deflected 70° over these spanwise extents. For the wing leading edge section fras 0 = 0.546 to 1.00, the slat was the same as the fully swept case.

Figure 2(d) shows the leading edge configura:ions used during the second wind tunnel investigation. Highly cambered slats with increased chords of 0.3397 m (1.114 ft) and 0.3086 m (1 -013 ft) wece instailed at the unswept leading-edge sections. These slats could be deflected either 60° or 70' with a 0.015~ gap. In addition, these slats were used as Krueger flaps that co~id be deflected either 70° or 80°. For the wing leading edge section from n = 0.546 to 1.G0, the slat used during the first investigation was modified to give more camber at its trailing edge and was relocated to give a 0.01% gap with respect to the modified wing leading edge. A slat deflection of 6S0 was used during the iavestigation.

The leading edge configurations used during the investigations are summarized in Table 11.

Leading-edge BLC system

Figure 2(e) shows the leading-edge BLC sysrem and nozzle arrangement used during both wind tunnel investigations.

Air fcr the blor~ing BLC nozzles was supplied by a centrifugal compressor located at the forward portion of the fuselage. This compressor was driven by tw variable frequency 300 horsepower electric motors coupled together.

Ihe air from the compressor outlet was ducted as shown with appropriate valvii~g to the wing leading-edge BU; nozzles and aileron BLC nozzles.

For the first investigation, the leading edge BLC nozzle was located between the fuselage and the outboard nacelle at 0.0075~from the swept leading edge kith a gap of either 0.318 cm or 0.160 cm. An air pressure ratio that ranged from 1.17 to 1.41 was used during this investigation.

For the second wind tunnel investigation, BLC nozzles were installed at the unswept leading edge sections, both sides of the inboard nacelle, and the inboard side of the outboard nacelle. The leading-edge BU: nozzle was located SSO from the wing chord plane with a gap of 0.101 cm. The nacelle BK nozzle was located lSO from the vertical reference line and intersected the wing leading-edge BU: nozzle. The nacelle nozzle had a length of 20.32 cm and a gap of 0.203 cm. An air pressure ratio that ranged from 1.17 to 1.33 was used at both nozzles during the second investigation.

Trailing-edge flap system

A Coanda plate surface was installed over the double-slotted flap from I, ;. 0.11 to 0.70 as shown in figure 2(f). The flap was the same as reported in reference 1 except for the itrcreased spanwise extent of the Coanda surfpce. Separate Coanda plates were used to provide a jet flap deflection (6f) of 30° and 7S0 measured between the flap trailing edge and the wing chord plane. For 6f = 93O a 0.254 m (0.834 ft) chord extension was added at the trailing edge of the Coanda plate used for 6f = 7S0.

Aileron

As shown in figure 2(g) a 0.35~plain aileron with BLC extended from 0 = 0.70 to 1.0 and could be deflected from O0 to 23O measured perpendicular to the hinge line. For the first wind tunnel investigation, the BLC rmzzle was located 30° ahead of the 0.65~line. For the second wind tunnel in- vestigation the nozzle was relocated lSO ahead of the 0.6% line to improve the air flow over the aileron radius. A nozzle gap of 0.089 cm with a pressure ratio that ranged from 1.16 to 1.39 was used during the investiga- '-ions.

Propulsion

The upper surface blowing flap and nozzle arrangement is shown in figure 2(h). Tine JTlSD-1 engines were used during the investigations and were housed in nacelles as shown in the figure. The erigines have a bypass ratio of 3 and a nonaal maximum gross thrust of 2200 pounds. The engine ccnterline was coincident with the nacelle centerline and was pitched up lo with respect to the wing chord plane. The inboard and outboard engine centerlines were located at n = 0.256 and 0.480, respectively.

The engine nozzle configuration used during both wind tunnel investiga- tions is shown in figure 2(h). The nozzle had an aspect ratio of 5.5 and corresponded to nozzle D of reference 1.

During the investigations, two vanes were located on each side of the nacelles close to the wing leading edge as slown in figure 2(h). These were instailed to generate a vortex to improve the flow along the side of the nacelle and the wing upper surface. In addition, a wing fence was installed during the investigations at TI = 0.37 as shown ia figure 2(h) to decrezse the exhaust flow ifiteraction between the inboard and outboard engines. Vortex generators were zlso installed briefly on the wing upper surface adjacent to the inboard side of the outboard nacelle as shown in the figure. The nacelle contours used during the investig~tionsare defined in figure 2(i). During the first wind tu~elinlpestigation the lower half of the inboard and outboard nacelle cross sections were modified to elliptical sections from station 2 to 7 as shown in the figure. This was done to improve the upflow over the wing leading ecige.

Tail

The geometry of the horizontal and vertical tails is shown in figure 2(a). These tails are the same ones used in reference 1. The horizontal tail detail is shown in figure 2(j). The horizontal tail i-.cidence and were set at 0" when the tail was installed. The vzrtical tail was on the model throughout both investigations.

CORRECTIONS

The data were corrected for wind-tunnel wall constraints. These cor- rections were determined by considering only the aerodynamic lift of the model (Ci) that resulted after the jet reaction components had been subtracted from the data as follows:

C; = CL - nf CJ (sin (6 + j a,))

a = R + .4175 C; u

Cm = C + 0.025 C; (horizontal tail on tests only) mu The engine thrust values defining CJ were based on the calibration of the engine static thrust variation with engine fan rotational speed. The calibration of each engine was obtained from wind tunnel scale measurentents with the flap undeflected. The 6j and nf values used in the corrections are shown in figure 3. These values were obtained in the wind tunnel with four engines operating and with the wind off. Evaluated from tunnel balance measurements, is the resultant thrust (FR) divided by stztic thrust (Fg) . The data that are presented in this report are not corrected for ram drag, but for reference the variation of ram drag with CJ is presented in figure 4. TESTING AND PROCEWRE

The data to compute static jet turning angle and resultant thrust with the flap deflected were recorded in five second intervals during the period the four engines were accelerated simultaneously from idle setting to a thrust setting of 1000 pounds per engine. This was done to obtain data before the engine thrust generated airflow in the test section. The tunnel overhead doors were apened when these data were recorded.

Forces and moments were measured through an angle-of-attack range of -8O to 28O. Tests were conducted at Reynolds numbers of 2.1 x lo6 and 3.0 x lo6 corresponding to dynamic pressures of 239 and 479 ~/m~(5.0 and 10.0 psf) , respectively, and based on a mean aerodynamic chord of 1.69 m (5.56 ft).

Tests With Constant CJ and Varying Angle of Attack

Four engines operating- A constant CJ was maintained as angle of attack was varied for each flap configuration investigated. The nominal C.1" values used in most cases during the investigation are as follows:

CJ (4 engine) L, Wm2

The variables studied were jet flap deflection, leading-edge BLC, nacelle BLC, wing leading-edge inboard sweep, and leading-edge slat or flap deflection. Tests were conduct4 with and without the t ~izontaltail.

Three engines operating- Tests were conducted with either the left hand outboard or inboard engine out at 6f = 30' and 90°. In addition, tests were conducted with the right hand outboard engine out at 6f = 90°. In most cases, the Coanda surfsce behind the inoperative engine was left on.

Tests With Consta~tCJ and Varying Angle of Sideslip

A constant CJ was maintained at 4, = 4O as 8 was varied from 8O ta -lgO for most cases. Tests were conducted with all engines operating, or left 'hand outboard engine out. RESULTS AND DISCUSSION

The static turning efficiencies (nf) and static turning angles (tij) are shown in figure 3. The variation of C with Cj is shown in figure 4. Dram The jet exhaust total pressure distributionsbehind the ezlgine nozzle and at the Zlap trailing edge (see figure 2(h)) along the inboard and outboard engine centerline are shown in figure 5. The basic serdynamic data are presented in figures 6 through 32. An index to these data is given in Table 111. The flap chordwise surface pressures at several spanwise stations are shown in figure 33. The variation of average downwash angle with angle- of-attack at the horizontal tail location is shown in figuzes 34(a) and (b) for 15~= 30° and 90°, respectively. These data were obtained from a doun- wash rake mounted at the tail location as shown in figure 2(a). The variation of CL with CPLE with the swept and unswept leading edge is shown in figure 35. A comparison of C and ac values between the tvo engine model of 'max lnax reference 1 knd tfie four engine model with the wing fully swept is shown in figure 36. The variation of Cy, Cn, an3 C;! with Cj at 4, = 4' is shown in figures 37(a) and (b) for 6f = 30' and 90°, respectively with either the inboard or outboard engine out case.

Static Turning

The 6j and rlf values shown in figure 3 were obtained with four engines operating at equal thrust. The engine nozzle which was used during the in- vestigations corresponded to nozzle D of reference 1. A comparison with the results of reference 1 is also shown in the figure. Slightly higher values of 6j and of were obtained with four engines operating when the results are compared with one engine operating of reference 1. However, the result is nearly the same between two engine operation of reference 1 and the four engine operation. As mentioned in reference 1, higher 6 value was obtained with multi-engine ope~ation. This was probably due to t i e jet exhaust spreading over the top of the fuselage with one engine operating.

Improvement of Maximum Lift and Stall Angle

Reference 3 discusses the problem and the subsequent improvement of maximum lift and stall angle of the model in greater detail. As inaicated in the reference, the large nacelles extending well above the wing upper surface caused high upwash angles between the nacelles and between the inboard nacelle and the fuselage. This created an adverse pressure gradient at the leading edge and Zed to flow separation in these areas which affected maximurn lift and stall angle. The deterioration of these values when the outboard engines were installed is shown in figure 36 in a comparison between the two engine model of reference 1 and the four ergine model. As mentioned previously, these models were nearly identical except for the number of engines. The values of C and the stall angle &re lowered approximately %ax by 1.0 and 8' to 120, respectively, from CJ = O to 2.9. Efforts were made to improve maximum lift and stall angle by changing the leading edge and nacelle configuration at the critical areas. Figure 13 shows the effect of modifying the nacelle contour (see figure 2(i)) near the wing leading edge. Maximum lift is improved slightly, but the stall angle remained the same. The effects of leading-edge BLC on the swept leading edge ald unsweeping the leading edge near the critical areas are also shown in the figure. In either case, C and a~ increased (approximately 10 percent =maX Is..ur and 5.S0, respectively) over that without any treatment on the swept leading edge. Additional improvement was obtained with the inboard leading-edge un- swept by applying blowing along the nacelle sides as shown in figure 14. cbax and CIC Lmax values increased 4 percent aad 7O, respectively. The addition of leading edge BLC to the unswept leading edge sections along with nacelle blowing did not give further improvement as shown in the same figure.

The effect of slat and Krueger flap deflections at the unswept leading edge sections on C is shown in figure 16. The higher slat deflection or Lmax Kmeger flap deflection did not provide any significant improvement in C Lmax' Neither thz combination of nacelle vanes and wing nor the combination of nacelle vanes and wing fence or nacelle vanes alone provided any sizeable maximum lift i.mprovements as shown in figures 13 and 19, respective1)-.

Longitudinal and Lateral Characteristics With an ingine Out at Zero Sideslip

The effects of engine out are shown in figures 23, 26, and 27. As shown in these figures, higher values of lift were obtained with the outSoard engine out compared to the inboard e~rgineout case, but the values of drag remained essentially the same for either case. As expected, the outboard engine out case provided a greater nose up pitching moment.

The variation of Cy, Cn, and Ct with CJ with either the outboard or the inboard engine out on the left hand side are shown in figures 37(a) and (b) for 6f = 30° and 90': respectively. As shown in the figures, the values of rolling moment with the outboard engine out was approximately twice those with the inboard engine out for either 6? = 30° and 90'.

REFERENCES

1. Aoyagj , Kiyoshi; Falarski, Michael D. ; and Koenig, David G. : Wind Tunnel Investigation of a Large-Scale Upper Surface Blown-Flap Transport Model Having Two Engines. NASA TM X-62,296, 1973. 2. Falarski, Nichael D.; Aoyagi, Kiyoshi; and Koenig, David G.: Acoustic Characteristics of a Large-Scale Wind-Tunnel Model of an Upper-Surface Blown Flap TransLwrt Havi~gTwo Engines. NASA TM X-62,319, 1973.

3. Koenig, David G.; and Aoyagi, Kiyoshi: Maximust Lift of Upper Surface Biowing STOL Aircraft with Swept . AIAA Paper 75-868, June, 1975. TABLE I- WDIFIED ORDINA- (-ISE DIRECTION) TIP SECTION

I I

iI

TAB1.E 111- LIST 01: RASIC IIATA PI(;IIRliS - CONI'INUKD

LE BLX effect with swept Inbound LE and effect offI of unswept inboard LE - Unswupt inboard LE

Effect of nacelle BIZ Ill I with and without 1.E BLC mGt of unswept LE devices .. Effort of aileron deflection Effect of LE BU:

mcrnacelle vanes and wing fenca a. cOC 0 c 0 -a- men- -4- .) e:. a%-,-- .- ua m o 3

-- iZfZ 2: 2-2 -9 2-2 mar I. 8.8 2%*;I a f+!%Z+* -2 -7: - --+ ---7- -. DC o ceo o~c.cceoo~oo, YmU-OdLI-, , J i -0..7LY1-, I m 10 m oe o rs_s*Iocl -- Ic I') s A *-0 *-.-* u Ci

(d) Leading-edge slat arrangement used during Test 441.

Figure 2.- Continued.

/ Ma%? +'- I. A'L ~~0.mm..r1~,~)-.u MJZD. LILLCLCTIDV~~~Lb8Lwr.t-.

h RSC .PLY I

(WHEN CMNAM RATES RLWDI

(f) Trailing-edge flap arrangement. Figure 2. - Continued.

Figure 3.- Flap static turning efficiency and turning angle. Figure 4.- Variation of CD with CJ. ram

00'6 00'9 OC'L 00.9 OJ'S OO'h 00'6 0 00'1 13 . . L ti -..I 0 ce w. are, 0" ; -4 C, arn- =: +- 0 st 0 e---v C, 0 .j Z1 L mn 3 &4 M a =- -.-I s 2% uac U 0 4 0 (2 -. CO W 40 -;1 OQI 7 vu c. -4 0 M ru cw 0 0 C -II .- * -u > re .v 0 t--0 - da3 4 -0 0'44 U st@ a 3a C M

GO'01 30'6 00'8 00.L 00.9 GO'S 06.R 00.c 00.2 00.1 73

0

00'0t 00'6 OO'B OO'L 00'9 00'5 OO'h OO't 00'2 OO'L D 13 0 9

s0 uU c Z

x =a

U QN C

S WM U U U -4 It W A* 4 (d mcu c .rlw - a-,-I ka- oC 3 we c, US Q) (do C

a0OD ca he 0 Q 1 do Q) 0 8-C -x P, N.k N J M Qa C k c -4

13 "0

10'0 00'4 00'0 00'9 OO'h OO'E OO'Z 00'1 73

03'6 00.0 DO'L 00': OC'S12 20 P m s 1 -L sir ;1

P

JO'CI -3'6 00'8 D3'L O.'I 00's 00'8 OC'C Oe.2 LO': 73

-1 6 -12 -V 0 Y j (BC?R) 4

Figure 29.- Variation of side force, yawing-moment, and rolling-moment coefficients with sideslip; = 30'. = 0°, CYLE = 0, = 0, a, = 4', wing fence off, nacelle vane off, P KAC LE configuration 6, tail on. Figure 30.- Variation of side force, yawing-moment, and rolling-moment coefficients with sideslip and with an engine out;

hail = 0, CpLE = 0, CpNAC = 0, aU = 4O, wing fence vane off, LE configuration 6, tail on. Figure 31.- Variation of side force, yzwing-moment, and rolling-moment coefficient with sideslip; 6f = 90°, dail = 0°, a, = 4=, wing fence off, nacelle vane off, LE config~~ation5, tail 011. (a) L.H. outboard engine out.

Figure 32.- Variation of side force, yawing-ndment, and rolling-moment coefficient with sideslip and with an engine out; 6f = 90°, 6,il = 0'. CpLE = 0, CusAC = 0, ,au = 4O, wing fence off, nacelle vane off, LE configuration 5, tail on. (b) L.H. inboarc! ^nc:.~coat.

Figure 32. - Concluded. Figure 33.- Flap surface pressures at sever-1 spandise statims, 6f = 90°, CJ = 3.0, q = 239.40 ~/m*, crvg. exhaust oressure ratio = 1.06. (b) a,: = 20°.

Figure 33. - Concluded. (a) = 30'.

Figure 34.- Variation of average downwash angle with angle-of-attack. fb) Ef = 9!1°.

F igure 24. - Concluded . Figure 35.- Variation of CL with Cp with the swept 2nd unswept inboard LE leading edge; = 23'. Figure 36.- Cgmpsri wn of CL and ;r~ between the two engine m3u Lmax ~ndthe four en:ine US3 model with the leading edge fully shtpt

:;.-I0; 'f = -50. (a) bf = 30'.

Figure 37.- Variation of C Cn, Cl with CJ; = 0°, 4, = 4O. Y' (b) fif = 90'.

Figure 37.- Concluded.