RESEARCH MEMORANDUM IN THE LANDING CONFIGURATION WIND-TUNNEL INVESTIGATION OF "HE LOW-SPEED STATIC AND

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1 RESEARCH MEMORANDUM WIND-TUNNEL INVESTIGATION OF "HE LOW-SPEED STATIC AND ROTARY STABILITY DERIVATIVES OF A 0.13-SCALE MODEL OF Tm D3UGLAS D-558-IC ALRPLANh IN THE LANDING CONFIGURATION

2 1v RESFARCH MEMORANDUM. WIND-TUNNEL INYESTIGATION OF THE LOW-SPEED STATIC AM3 ROTARY STABILITY DE UVA!I'IVES OF A 0. U-SCATX MODEL OF THE DOUGLAS D AIRPIJUG3 IN THE LANDIE CONFIGURATION By M. J. QueiJo and Evalyn G. Welb SUMMARY A wind-tunnel investigation has been made to determine the lowspeed static and rotary stability derivatives of a 0.13-scale model of the Douglas D airplane in the lending configuration. The Eft coefficient of the mdel varied Unearly with angle of attack up to a maximum lift coefficient of 1.24 which occurred at an angle of attack of U0. The Ufi-curve slope was about 0.06 per degree in this range. The model was longitudinallystable in the angle-of-attack range from Oo to 16O, with a static margin of about 16 percent of the wing mean aerodynamic chord over most of this range. The model was approximately neutrally stable near an angle of attack of Il0. The directional stability of the model decreased slowly with increase in angle of attack up to an angle of attack of about l3o. At higher angles, the stability deteriorated more rapidly. The yawing moment due to rolling velocity was negative throughout the angle-ofattack range, and the magnitude of the tail co-ntribution to this moment near zero angle of attack indicated a stronger sidewash effect for the flapped wing than generally has been obtained for plain wings. The derivatives associated with yawing flow were nearly constant for angles of attack from Oo to about uo, but varied considerably at higher angles. INTRODUCTION Various investigations have shown that the dynamic lateral stability characteristics of high-speed aircraft are critically dependent on

3 2 NACA RM L52G07 certain mass and aerodynamic parameters. and, hence, that reliable estimates ofthe dynamic stability of such aircraft catl be made only if these parameters are determined accurately. The static derivatives of an airplane can be determined accurately by 11~3811s of' conventional wind-tunnel tests of a model; however, only a few facilities are available for measuring rotary (rolling and m~ing) derivatives. The Langley stability tunnel, which is equipped with facilities for simulating rolling and yawing flaw, was qtilized to make available measured lowspeed static and-rotary derivatives of a mdel of the Douglas D airplane in the landing configuration (slats, flaps, and landing gear extended). The measured low-speed parameters of the same model with slats, flaps, and landing gear retracted are given in reference 1. SYMBOLS AND COEFFICIENTS The data presented herein are in the form of standard IUCA coefficients of forces and moments which are referred to the system of stability axes (fig. 1) with the origin.at the projectiun of the quarterrchard..point of the wing mean aerodysamlc cbord on the plane of symmetry. This syeltem of axes is defined as an orthogonal system having the origin at the center of gravity and in which the Z-axis is in the plane of symmetry and perpendicular to the relative wind, the X-axis i8 in the plane of symmetryand perpendicular to the Z-axis, and the Y-axis is perpendicular to the phne of symetry. Positive directions of forces, moments, and displacements-are sham in figure 1. b C - C wing span, ft local wing chord, parallel to plane mean aerodynamic chord, ft of symnetry, ft P rolling angular velocity, radians/sec 9 dynamic pressure, $, lb/sq ft 2 r yawfng angular velocity, raaans/aec S V a P wing area, sq ft free-stream velocity, ft/sec angle of attack, deg sides angle, lip radians 4

4 . Y angle of clirib, deg 3 D. 4f angle of yawy deg # angle of roll, deg P D L Y M N mass density of air, slugs/cu ft drag, lb lift, lb side force, lb pitching moment, ft-lb yawing moment, ft-lb 2 rolling moment, ft-lb CD drag coefficient, D/qS CL lift coefficient, L/qs CY side-force coefficient, Y/QS cm pitching-mment coefficient, M/qS Cn yawing-mment coefficient, N/q% Cl rolling-moment coefficient, Z/qSb - c - &Y yp - 2v L

5 4 NACA RM L52GO7. cyr = -&Y rb agv czr = kl - rb APPARATUS, MODEL, AND "S The tests of the present investigation were conducted in the 6-foot-aameter rolling-flow and 6- by 6-foot yawing-flow test sections of the Langley stability tunnel, in which rolling and yawing flow are simulated by curving the air stream about a stationary mdel (refs. 2 and 3). 'A single-strut support was used to attach the mael to a sixcomponent balance sys tem. The model used in the investigation was a 0.13-scale~model of the Douglas D-5g-11 airplane and was constructed of laminated mahogany. A drawing of the model Is given as fiwre 2, with details of the flaps and slats given in figure 3. Pertinent geometric characteristics of the model are listed in table I, and a photograph of the model used in the investigation is presented as figure 4. Tests in straight and rolling flow were made at a dynamic pressure of 39.7 pounds per-square foot, which corresponds to a Machnumber of 0.17, and a Reynolds nmber of 1,100,000 based on the wing mean aerodynamic chord. The tests in sideslip and in y-awing flow were made at a dyndc pressure of 24,9 pounds per square foot, corresponding to a Machnumber of 0.13 and a Reynolds number of 865,000. Tests were made with the complete mdel and also with the wing-fuselage combination.

6 NACA m ~ 52~07 1: 5 CORRECTIONS Approximate corrections for jet-boundary effects were applied to the angle of attack by the methods of reference 4 and to the pitchingmoment coefficfent by the methods of reference 5. RFSULZS AND DISCUSSION Static Longftudinal Characteristics J- The lift coefficient of the model in the landing configuration increased linearly with angle of attack up to CL = 130t and the liftcurve slope &I,/& was about 0.06 per degree (fig. 5). A maximum lift coefficient of 1.24 was attained at a. = -U0, and remained near that value for angles of attack from l3o to 22O. The model was lonetudinally stable (negative &,&a) in the angle-of-attack range from ' 0 to about 16O with a static margin of about 0.l6F over most of the range. The mdel was approximately neutrally stable near a = llo. At angles of attack above about 16O, the pitching-moment coefficient changed erratically with angle of attack. Static Lateral Characteristics The directional instability of the wing-fuselage combination (negative Cnp) was approximately constant through the angle-of-attack rmge (fig. 6). Addition of the tail surfaces made the model directionally stable throughout most of the angle-of-attack range; however, the degree of stability generally decreased with increase in angle of attack. The effective-dihedral parameter C.lp was approximately the same for the wing-fuselage combination as it was for the complete mdel, and generally increased negatively with an increase in angle of attack. Characteristics in Rolling Flow The aerodynamic derivatives of the model in simulated roll are shown in figure 7 a5 curves of Cyp, Cnp, and Czp plotted against. angle of attack. Addition of the tail surfaces to the wing-fuselage combination produced a negative increment of Cnp. From geometric considerations, such as those of reference 6, a positive increment to would have been expected near a = Oo from addition of the tail cnp

7 '6 surfaces. The negative increment actually obtained probably is caused by changes in flow angularity at the tall. plane, associated wfth wing wake characteristics (ref. 7). It appears that deflection of the flaps has a rather powerful effect. on the flow angularity at the. tazl since, in investigations with models having plain wings, the positive tail contribution to Cnp at a = Oo, although reduced by wing wake angularity, has not been made negative (for example, 6ee refs. 7, 8, or 9). Characteristics in Yawing Flow The yawing-flow parameters Cyr, Cnr, and Cz are plotted against r angle of' atkack in figure 8. These parameters remained approximately constant for angles of attack up to about. U0 for the model with the tail surfaces on or off. At angles of attack greater than 13', the pkrameters varied over a rather large range with increase in angle of attack. CONC WSIONS An investigation was made in the Langley stability tunnel to determine the low-speed static and rotary stability derivative8 of a 0.13-scale model of the Douglas D airplane in the-laiding configuration. The results of the Fnveatigation have led to the following conclusions: 1. The lift coefficient varied linearly with angle of attack up to a maximum lift coefficient of 1.24, which occurred at an angle of 8ttaCk of 13'. The lift-curve slope was about 0.06 per degree in this range. 2. The model had static lo itudinal stability in the angle-of- attack range from Oo to about 16 3 with a static margin of about 16 per- cent of the wing mean aerodynamic chord over most of the range. The model was approximately neutrally stable near an angle of attack of 1l0. 3. The directional stability of the mdel decreased slowly with increase in aagle of attack up to about 13'. At higher angles of attack, the directional stability generally decreased more rapidly. 4. The yawing moment due to roll C was negative throughout the np angle-of-attack range. The magnitude of the tail contribution to Cn P at low angles of attack indicated a stronger sidewash effect in roll for the flapped wing than generally hag beep obtained for plain wings.

8 NACA RM L5xO The.derivatives associated with- yawing flow were about constant in the angle-of-attack range from Oo to U0, and varied considerably with angle of attack above Uo. Langley Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field, Va.

9 8 NACA RM L52GO7 REFERENCES 1. QueiJo, M. J., and Goodman, Alex: Calculations of the DynWc Lateral Stability Characteristics of the Douglas D Airplane in High-Speed-..Flight. for.various Wing had.&& and-alt itudes.- NACA FM ~50~l6a, MacLachlan, Robert, and Letko, Wiliam: Correlation of Two Experimental Methods of Determining the Rolling Characteristics of Unswept Wings. NACA TN 1309, Bird, John D., Jaquet, Byron M., and Caran, John W.: Effect of Fuselage and Tail Surfaces on Low-SpeedYawing Characteristics of a Swept-WingModel AS Determined in.curved-flow Test Section of Langley StKbility Tunnel. NACA Tw 2483, 191. (Supersedes NACA RM L8G13.) 4. Silverstein, Abe, and White, James A.: Wind-Tunnel Interference With Particular Reference to Off-Center Positione of the Wing and to the Downwash at the Tail. NACA Rep. 547, Gillis, Clarence L., Polhamus, Edward C., and Grey, Joseph L., Jr. : Charts for Determining Jet-Boundary Corrections for Complete Models in 7- by 10-Foot Closed Rectangular Wind Tunnels. WCA L5G31, Bamber, Milled J. : Effect af Some Present-Day Airplane Design Trends on Requirements for Lateral Stability. NACA TN 814, Michael, Wiliam H., 3r. : Analysis of the Effects of Wing Interference on the Tail Contributfon to the RollFng Derivattvee. TM 2332, "- NACA " ,." 8. Letko, Wiliam, and Riley, Donald R. : Effect of an Unswept Wing on the Contribution of Unswept-Tail Configurations to the Low-Speed Static- and Railing-Stability Derivatives of a Midwing Airplane Model. NACA TN 2175, Qolhart, Walter D.: Influence of Wing and Fuselage on the Vertical- Tail Contribution to the Law-Speed Rolling Derivatives of Midwing Airplane Models With 45' Sweptback Surfaces. NACA TN 2587, 1951.

10 2v.. NACA RM L52GO7 wing : TABLE I.. DIMENSIONS AND CEARACTERISTICS OF- HODEL Root airfoil section (normal to 0.33-chord lwe)... Tip airfoil section (normal to 0.33-chord line)... Total area, sq in... Span. in.... Mean aerodynamic chord, in.... Root chord (parallel to plane of symmetry), in.... Tip chord (parallel to plane of symmetry), in... Taper ratio... Aspect ratio Sweep at 0.33-chord line, deg... Incidence, deg... Dihedral, deg... Total flap area, sq in... Horizontal Tail: Airfoil section (no&l to 0.35-chord line)... NACA Total area. sq in span. in Mean aerodynamic chord. in Root chord (parallel to plane of symmetry). in Tip chord (parallel to plane of symmetry) in Taper ratio Aspect ratio Sweep at 0.35-chord line. deg Incidence (from fuselage center line). deg Tail length (from F/4 of wing to F/4 of tail). in Tail height ( from fuselage center iine) in NACA NACA Vertical Tail : Airfoil section (normal to 0.45-chola line)... NACA Root chord (parallel to fuselage center line). in Height. from f'uselage center line. in Sweep - at Q.@-chord line. deg Fuselage : Length. in Maximum diameter. in Fineness ratio

11 10 NACA RM ~52~07 Figure 1.- System of atability axee. Arrows indicate positive direction of forces, moments, and diaplacement8.

12 .jl I b Figure 2.- Model uaed In the inveetl&atlon. All dlmensims glven In Inche.8.." P...

13 12 - MACA RM L52GO7 7jpcd sechn through fhp norma/ tu & Figure 3.- Details of slats and plain flaps. All dimeneions given in Inchee.

14 I r I h I

15 14./ 0 -.I /2 / Anqle of ufhck, as, de9 Fi. w e 5.- Variation of CL, CD, and C, with a for 8 0.l3-scale model ofthe D airplane in the landing configuration. Mach number, 0.17; Reynolds number, 1,100,000.

16 NACA RM L5xO L.I /2 / Angle of afhck, a3, deg Figure 6. - Variation of Cyp, Cnp, and C with a Tor a 0.13-scale 28 model of the D-3S-11 airplane in the landing configuration. Mach number, 0.17; Reynolds number, 1,100,000.

17 11 c - NACA RM L52GO7 G 2P Angle of aftuch, a;, de9 Figure 7.- Variation of Cyp, C and C2 with a for a O.13-sale np' P model of the D airplane in the landing canffguration. Mach number, 0.17; Reynold6 number, 1,100,000.

18 3v a L / Angle of attack, zl deg Figure 8. - Variation of Cy Cnr and C 2, with a for a 0.13-scale rj model of the D airplane In the landing configuration. Mach number, 0.13; Reynolds nuniber, 863,000. fr " -

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